National transonic facility Mach number system
Kern, F. A.; Knight, C. W.; Zasimowich, R. F.
1985-01-01
The Mach number system for the Langley Research Center's National Transonic Facility was designed to measure pressures to determine Mach number to within + or - 0.002. Nine calibration laboratory type fused quartz gages, four different range gages for the total pressure measurement, and five different range gages for the static pressure measurement were used to satisfy the accuracy requirement over the 103,000-890,000 Pa total pressure range of the tunnel. The system which has been in operation for over 1 year is controlled by a programmable data process controller to select, through the operation of solenoid valves, the proper range fused quartz gage to maximize the measurement accuracy. The pressure gage's analog outputs are digitized by the process controller and transmitted to the main computer for Mach number computation. An automatic two-point on-line calibration of the nine quartz gages is provided using a high accuracy mercury manometer.
Allan Brian G.; Owens, Lewis, R.
2006-01-01
This paper will investigate the validation of a NASA developed, Reynolds-averaged Navier-Stokes (RANS) flow solver, OVERFLOW, for a boundary-layer-ingesting (BLI) offset (S-shaped) inlet in transonic flow with passive and active flow control devices as well as the baseline case. Numerical simulations are compared to wind tunnel results of a BLI inlet conducted at the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. Comparisons of inlet flow distortion, pressure recovery, and inlet wall pressures are performed. The numerical simulations are compared to the BLI inlet data at a freestream Mach number of 0.85 and a Reynolds number of approximately 2 million based on the length of the fan-face diameter. The numerical simulations with and without wind tunnel walls are performed, quantifying effects of the tunnel walls on the BLI inlet flow measurements. The wind tunnel test evaluated several different combinations of jet locations and mass flow rates as well as a vortex generator (VG) vane case. The numerical simulations will be performed on a single jet configuration for varying actuator mass flow rates at a fix inlet mass flow condition. Validation of the numerical simulations for the VG vane case will also be performed for varying inlet mass flow rates. Overall, the numerical simulations were able to predict the baseline circumferential flow distortion, DPCPavg, very well for comparisons made within the designed operating range of the BLI inlet. However the CFD simulations did predict a total pressure recovery that was 0.01 lower than the experiment. Numerical simulations of the baseline inlet flow also showed good agreement with the experimental inlet centerline surface pressures. The vane case showed that the CFD predicted the correct trends in the circumferential distortion for varying inlet mass flow but had a distortion level that was nearly twice as large as the experiment. Comparison to circumferential distortion measurements for a 15 deg clocked 40 probe
Shrewsbury, George D.; Vadyak, Joseph; Schuster, David M.; Smith, Marilyn J.
1989-01-01
A computer analysis was developed for calculating steady (or unsteady) three-dimensional aircraft component flow fields. This algorithm, called ENS3D, can compute the flow field for the following configurations: diffuser duct/thrust nozzle, isolated wing, isolated fuselage, wing/fuselage with or without integrated inlet and exhaust, nacelle/inlet, nacelle (fuselage) afterbody/exhaust jet, complete transport engine installation, and multicomponent configurations using zonal grid generation technique. Solutions can be obtained for subsonic, transonic, or hypersonic freestream speeds. The algorithm can solve either the Euler equations for inviscid flow, the thin shear layer Navier-Stokes equations for viscous flow, or the full Navier-Stokes equations for viscous flow. The flow field solution is determined on a body-fitted computational grid. A fully-implicit alternating direction implicit method is employed for the solution of the finite difference equations. For viscous computations, either a two layer eddy-viscosity turbulence model or the k-epsilon two equation transport model can be used to achieve mathematical closure.
Powers, Sheryll Goecke; Huffman, Jarrett K.; Fox, Charles H., Jr.
1986-01-01
The effectiveness of a trailing disk, or trapped vortex concept, in reducing the base drag of a large body of revolution was studied from measurements made both in flight and in a wind tunnel. Pressure data obtained for the flight experiment, and both pressure and force balance data were obtained for the wind tunnel experiment. The flight test also included data obtained from a hemispherical base. The experiment demonstrated the significant base drag reduction capability of the trailing disk to Mach 0.93 and to Reynolds numbers up to 80 times greater than for earlier studies. For the trailing disk data from the flight experiment, the maximum decrease in base drag ranged form 0.08 to 0.07 as Mach number increased from 0.70 to 0.93. Aircraft angles of attack ranged from 3.9 to 6.6 deg for the flight data. For the trailing disk data from the wind tunnel experiment, the maximum decrease in base and total drag ranged from 0.08 to 0.05 for the approximately 0 deg angle of attack data as Mach number increased from 0.30 to 0.82.
Drake, Hubert M; Mclaughlin, Milton D; Goodman, Harold R
1948-01-01
Results are presented of tests up to a Mach number of 0.92 at altitudes around 30,000 feet. The data obtained show that the airplane can be flown to this Mach number above 30,000 feet. Longitudinal trim changes have been experienced but the forces involved have been small. The elevator effectiveness decreased about one-half with increase of Mach number from 0.70 to 0.87. Buffeting has been experienced in level flight but it has been mild and the associated tail loads have been small. No aileron buzz or other flutter phenomena have been noted.
An Investigation of Transonic Resonance in a Mach 2.2 Round Convergent-Divergent Nozzle
Dippold, Vance F., III; Zaman, Khairul B. M. Q.
2015-01-01
Hot-wire and acoustic measurements were taken for a round convergent nozzle and a round convergent-divergent (C-D) nozzle at a jet Mach number of 0.61. The C-D nozzle had a design Mach number of 2.2. Compared to the convergent nozzle jet flow, the Mach 2.2 nozzle jet flow produced excess broadband noise (EBBN). It also produced a transonic resonance tone at 1200 Herz. Computational simulations were performed for both nozzle flows. A steady Reynolds-Averaged Navier-Stokes simulation was performed for the convergent nozzle jet flow. For the Mach 2.2 nozzle flow, a steady RANS simulation, an unsteady RANS (URANS) simulation, and an unsteady Detached Eddy Simulation (DES) were performed. The RANS simulation of the convergent nozzle showed good agreement with the hot-wire velocity and turbulence measurements, though the decay of the potential core was over-predicted. The RANS simulation of the Mach 2.2 nozzle showed poor agreement with the experimental data, and more closely resembled an ideally-expanded jet. The URANS simulation also showed qualitative agreement with the hot-wire data, but predicted a transonic resonance at 1145 Herz. The DES showed good agreement with the hot-wire velocity and turbulence data. The DES also produced a transonic tone at 1135 Herz. The DES solution showed that the destabilization of the shock-induced separation region inside the nozzle produced increased levels of turbulence intensity. This is likely the source of the EBBN.
A new numerical solver for flows at various Mach numbers
Miczek, F; Edelmann, P V F
2014-01-01
Many problems in stellar astrophysics feature low Mach number flows. However, conventional compressible hydrodynamics schemes frequently used in the field have been developed for the transonic regime and exhibit excessive numerical dissipation for these flows. While schemes were proposed that solve hydrodynamics strictly in the low Mach regime and thus restrict their applicability, we aim at developing a scheme that correctly operates in a wide range of Mach numbers. Based on an analysis of the asymptotic behavior of the Euler equations in the low Mach limit we propose a novel scheme that is able to maintain a low Mach number flow setup while retaining all effects of compressibility. This is achieved by a suitable modification of the well-known Roe solver. Numerical tests demonstrate the capability of this new scheme to reproduce slow flow structures even in moderate numerical resolution. Our scheme provides a promising approach to a consistent multidimensional hydrodynamical treatment of astrophysical low Ma...
Quasiperpendicular high Mach number Shocks
Sulaiman, A H; Dougherty, M K; Burgess, D; Fujimoto, M; Hospodarsky, G B
2015-01-01
Shock waves exist throughout the universe and are fundamental to understanding the nature of collisionless plasmas. Reformation is a process, driven by microphysics, which typically occurs at high Mach number supercritical shocks. While ongoing studies have investigated this process extensively both theoretically and via simulations, their observations remain few and far between. In this letter we present a study of very high Mach number shocks in a parameter space that has been poorly explored and we identify reformation using in situ magnetic field observations from the Cassini spacecraft at 10 AU. This has given us an insight into quasi-perpendicular shocks across two orders of magnitude in Alfven Mach number (MA) which could potentially bridge the gap between modest terrestrial shocks and more exotic astrophysical shocks. For the first time, we show evidence for cyclic reformation controlled by specular ion reflection occurring at the predicted timescale of ~0.3 {\\tau}c, where {\\tau}c is the ion gyroperio...
Kuhlman, John M.; Liaw, Paul; Cerney, Michael J.
1988-01-01
A numerical design study was conducted to assess the drag reduction potential of winglets installed on a series of low aspect ratio wings at a design point of M=0.8, C sub L=0.3. Wing-winglet and wing-alone design geometries were obtained for wings of aspect ratios between 1.75 and 2.67, having leading edge sweep angles between 45 and 60 deg. Winglet length was fixed at 15% of wing semispan. To assess the relative performance between wing-winglet and wing-alone configurations, the PPW nonlinear extended small disturbance potential flow code was utilized. This model has proven to yield plausible transonic flow field simulations for the series of low aspect ratio configurations selected. Predicted decreases in pressure drag coefficient for the wing-winglet configurations relative to the corresponding wing-alone planform are about 15% at the design point. Predicted decreases in wing-winglet total drag coefficient are about 12%, relative to the corresponding wing-alone design. Longer winglets (25% of the wing semispan) yielded decreases in the pressure drag of up to 22% and total drag of up to 16.4%. These predicted drag coefficient reductions are comparable to reductions already demonstrated by actual winglet designs installed on higher aspect ratio transport type aircraft.
Quasiperpendicular High Mach Number Shocks
Sulaiman, A. H.; Masters, A.; Dougherty, M. K.; Burgess, D.; Fujimoto, M.; Hospodarsky, G. B.
2015-09-01
Shock waves exist throughout the Universe and are fundamental to understanding the nature of collisionless plasmas. Reformation is a process, driven by microphysics, which typically occurs at high Mach number supercritical shocks. While ongoing studies have investigated this process extensively both theoretically and via simulations, their observations remain few and far between. In this Letter we present a study of very high Mach number shocks in a parameter space that has been poorly explored and we identify reformation using in situ magnetic field observations from the Cassini spacecraft at 10 AU. This has given us an insight into quasiperpendicular shocks across 2 orders of magnitude in Alfvén Mach number (MA ) which could potentially bridge the gap between modest terrestrial shocks and more exotic astrophysical shocks. For the first time, we show evidence for cyclic reformation controlled by specular ion reflection occurring at the predicted time scale of ˜0.3 τc , where τc is the ion gyroperiod. In addition, we experimentally reveal the relationship between reformation and MA and focus on the magnetic structure of such shocks to further show that for the same MA , a reforming shock exhibits stronger magnetic field amplification than a shock that is not reforming.
Chaotic behaviour of high Mach number flows
Varvoglis, H.; Ghosh, S.
1985-01-01
The stability of the super-Alfvenic flow of a two-fluid plasma model with respect to the Mach number and the angle between the flow direction and the magnetic field is investigated. It is found that, in general, a large scale chaotic region develops around the initial equilibrium of the laminar flow when the Mach number exceeds a certain threshold value. After reaching a maximum the size of this region begins shrinking and goes to zero as the Mach number tends to infinity. As a result high Mach number flows in time independent astrophysical plasmas may lead to the formation of 'quasi-shocks' in the presence of little or no dissipation.
Low Mach Number Fluctuating Hydrodynamics for Electrolytes
Péraud, Jean-Philippe; Chaudhri, Anuj; Bell, John B; Donev, Aleksandar; Garcia, Alejandro L
2016-01-01
We formulate and study computationally the low Mach number fluctuating hydrodynamic equations for electrolyte solutions. We are interested in studying transport in mixtures of charged species at the mesoscale, down to scales below the Debye length, where thermal fluctuations have a significant impact on the dynamics. Continuing our previous work on fluctuating hydrodynamics of multicomponent mixtures of incompressible isothermal miscible liquids (A. Donev, et al., Physics of Fluids, 27, 3, 2015), we now include the effect of charged species using a quasielectrostatic approximation. Localized charges create an electric field, which in turn provides additional forcing in the mass and momentum equations. Our low Mach number formulation eliminates sound waves from the fully compressible formulation and leads to a more computationally efficient quasi-incompressible formulation. We demonstrate our ability to model saltwater (NaCl) solutions in both equilibrium and nonequilibrium settings. We show that our algorithm...
High Reynolds number test of a NACA 651-213, a equals 0.5 airfoil at transonic speeds
Burdges, K. P.; Blackwell, J. A., Jr.; Pounds, G. A.
1975-01-01
Wind-Tunnel tests were conducted in the Lockheed-Georgia Company's compressible flow facility to determine the transonic two-dimensional aerodynamic characteristics of a NACA 65 sub 1-213 a = 0.50 airfoil. The results are correlated with data obtained in the NASA-Langley 8-foot transonic pressure tunnel and the NAE high Reynolds number 15x60-inch two-dimensional test facility. The tests were conducted over a Mach number range from 0.60 to 0.80 and an angle of attack range from -1 deg to 8 deg. Reynolds numbers, based on the airfoil chord, were varied.
Low Mach number fluctuating hydrodynamics for electrolytes
Péraud, Jean-Philippe; Nonaka, Andy; Chaudhri, Anuj; Bell, John B.; Donev, Aleksandar; Garcia, Alejandro L.
2016-11-01
We formulate and study computationally the low Mach number fluctuating hydrodynamic equations for electrolyte solutions. We are interested in studying transport in mixtures of charged species at the mesoscale, down to scales below the Debye length, where thermal fluctuations have a significant impact on the dynamics. Continuing our previous work on fluctuating hydrodynamics of multicomponent mixtures of incompressible isothermal miscible liquids [A. Donev et al., Phys. Fluids 27, 037103 (2015), 10.1063/1.4913571], we now include the effect of charged species using a quasielectrostatic approximation. Localized charges create an electric field, which in turn provides additional forcing in the mass and momentum equations. Our low Mach number formulation eliminates sound waves from the fully compressible formulation and leads to a more computationally efficient quasi-incompressible formulation. We demonstrate our ability to model saltwater (NaCl) solutions in both equilibrium and nonequilibrium settings. We show that our algorithm is second order in the deterministic setting and for length scales much greater than the Debye length gives results consistent with an electroneutral approximation. In the stochastic setting, our model captures the predicted dynamics of equilibrium and nonequilibrium fluctuations. We also identify and model an instability that appears when diffusive mixing occurs in the presence of an applied electric field.
Design of a continuously variable Mach-number nozzle
Institute of Scientific and Technical Information of China (English)
郭善广; 王振国; 赵玉新
2015-01-01
A design method was developed to specify the profile of the continuously variable Mach-number nozzle for the supersonic wind tunnel. The controllable contour design technique was applied to obtaining the original nozzle profile, while other Mach- numbers were derived from the transformation of the original profile. A design scheme, covering a Mach-number range of 3.0
Kelly, Thomas C.
1961-01-01
Results have been obtained i n t h e Langley 8-foot transonic pressure tunnel at Mach numbers from 0.40 t o 1.20 for several configurations of the Scout vehicle and f o r a number of related models. Tests extended over an angle-of-attack range from about -10 degrees to 10 degrees at a Reynolds number per foot of about 3.8 x 10 sup 6.
Parametric investigation of single-expansion-ramp nozzles at Mach numbers from 0.60 to 1.20
Capone, Francis J.; Re, Richard J.; Bare, E. Ann
1992-01-01
An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of varying six nozzle geometric parameters on the internal and aeropropulsive performance characteristics of single-expansion-ramp nozzles. This investigation was conducted at Mach numbers from 0.60 to 1.20, nozzle pressure ratios from 1.5 to 12, and angles of attack of 0 deg +/- 6 deg. Maximum aeropropulsive performance at a particular Mach number was highly dependent on the operating nozzle pressure ratio. For example, as the nozzle upper ramp length or angle increased, some nozzles had higher performance at a Mach number of 0.90 because of the nozzle design pressure was the same as the operating pressure ratio. Thus, selection of the various nozzle geometric parameters should be based on the mission requirements of the aircraft. A combination of large upper ramp and large lower flap boattail angles produced greater nozzle drag coefficients at Mach number greater than 0.80, primarily from shock-induced separation on the lower flap of the nozzle. A static conditions, the convergent nozzle had high and nearly constant values of resultant thrust ratio over the entire range of nozzle pressure ratios tested. However, these nozzles had much lower aeropropulsive performance than the convergent-divergent nozzle at Mach number greater than 0.60.
Low Mach Number Fluctuating Hydrodynamics of Diffusively Mixing Fluids
Donev, A; Sun, Y; Fai, T; Garcia, A L; Bell, J B
2012-01-01
We formulate low Mach number fluctuating hydrodynamic equations appropriate for modeling diffusive mixing in isothermal mixtures of fluids with different density and transport coefficients. These equations eliminate the fast isentropic fluctuations in pressure associated with the propagation of sound waves by replacing the equation of state with a local thermodynamic constraint. We demonstrate that the low Mach number model preserves the spatio-temporal spectrum of the slower diffusive fluctuations. We develop a strictly conservative finite-volume spatial discretization of the low Mach number fluctuating equations in both two and three dimensions. We construct several explicit Runge-Kutta temporal integrators that strictly maintain the equation of state constraint. The resulting spatio-temporal discretization is second-order accurate deterministically and maintains fluctuation-dissipation balance in the linearized stochastic equations. We apply our algorithms to model the development of giant concentration fl...
Milholen, William E., II; Jones, Gregory S.; Chan, David T.; Goodliff, Scott L.; Anders, Scott G.; Melton, Latunia P.; Carter, Melissa B.; Allan, Brian G.; Capone, Francis J.
2013-01-01
A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.
Mathematical and numerical aspects of low mach number flows
Energy Technology Data Exchange (ETDEWEB)
Schochet, St.; Bresch, D.; Grenier, E.; Alazard, T.; Gordner, A.; Sankaran, V.; Massot, M.; Sery, R.; Pebay, P.; Lunch, O.; Mazhorova, O.; Turkel, O.E.; Faille, I.; Danchin, R.; Allain, O.; Birken, P.; Lafitte, O.; Kloczko, T.; Frick, W.; Bui, T.; Dellacherie, S.; Klein, R.; Roe, Ph.; Accary, G.; Braack, M.; Picano, F.; Cadiou, A.; Dinescu, C.; Lesage, A.C.; Wesseling, P.; Heuveline, V.; Jobelin, M.; Weisman, C.; Merkle, C.
2004-07-01
Low Mach number flows represent a significant part of the various flows encountered in geophysics, industry or every day life. Paradoxically, the mathematical analysis of the equations governing these flows is difficult and on the practical side, the research of numerical algorithms valid for all flow speeds is continuing to be a challenge. However, in the last decade, both from the theoretical and the numerical sides, significant progresses were made in the understanding and analysis of the equations governing these flows. This conference intends to provide an up-to-date inventory of recent mathematical and numerical results in the analysis of these flows by bringing together both mathematicians and numericists active in this area. In the framework of the conference, a numerical workshop is organized which proposes to compute several challenging low Mach number flows: liquid flow around non-cavitating and cavitating NACA0015 hydrofoil, natural convection with large temperature differences, free convection, free surface flow, vessel pressurization. This document brings together the descriptions of the test cases of the numerical workshop and the abstracts of the conference papers: A 3D high order finite volume method for the prediction of near-critical fluid flows (G. ACCARY, I. RASPO, P. BONTOUX, B. ZAPPOLI); low Mach number limit of the non-isentropic Navier-Stokes equations (T. ALAZARD); simulation of cavitation rolls past a forward step with a bubble model (O. ALLAIN, N. BLASKA, C. LECA); flux preconditioning methods and fire events (P. BIRKEN, A. MEISTER); an adaptive finite element solver for compressible flows: application to heat-driven cavity benchmarks in 2D and 3D (M. BRAACK); comparison of various implicit, explicit, centered and upwind schemes for the simulation of compressed flows on moving mesh (A. CADIOU, M. BUFFAT, L. Le PENVEN, C. Le RIBAULT); low Mach number limit for viscous compressible flows (R. DANCHIN); some Properties of the low Mach number
Mathematical and numerical aspects of low mach number flows
Energy Technology Data Exchange (ETDEWEB)
Schochet, St.; Bresch, D.; Grenier, E.; Alazard, T.; Gordner, A.; Sankaran, V.; Massot, M.; Sery, R.; Pebay, P.; Lunch, O.; Mazhorova, O.; Turkel, O.E.; Faille, I.; Danchin, R.; Allain, O.; Birken, P.; Lafitte, O.; Kloczko, T.; Frick, W.; Bui, T.; Dellacherie, S.; Klein, R.; Roe, Ph.; Accary, G.; Braack, M.; Picano, F.; Cadiou, A.; Dinescu, C.; Lesage, A.C.; Wesseling, P.; Heuveline, V.; Jobelin, M.; Weisman, C.; Merkle, C.
2004-07-01
Low Mach number flows represent a significant part of the various flows encountered in geophysics, industry or every day life. Paradoxically, the mathematical analysis of the equations governing these flows is difficult and on the practical side, the research of numerical algorithms valid for all flow speeds is continuing to be a challenge. However, in the last decade, both from the theoretical and the numerical sides, significant progresses were made in the understanding and analysis of the equations governing these flows. This conference intends to provide an up-to-date inventory of recent mathematical and numerical results in the analysis of these flows by bringing together both mathematicians and numericists active in this area. In the framework of the conference, a numerical workshop is organized which proposes to compute several challenging low Mach number flows: liquid flow around non-cavitating and cavitating NACA0015 hydrofoil, natural convection with large temperature differences, free convection, free surface flow, vessel pressurization. This document brings together the descriptions of the test cases of the numerical workshop and the abstracts of the conference papers: A 3D high order finite volume method for the prediction of near-critical fluid flows (G. ACCARY, I. RASPO, P. BONTOUX, B. ZAPPOLI); low Mach number limit of the non-isentropic Navier-Stokes equations (T. ALAZARD); simulation of cavitation rolls past a forward step with a bubble model (O. ALLAIN, N. BLASKA, C. LECA); flux preconditioning methods and fire events (P. BIRKEN, A. MEISTER); an adaptive finite element solver for compressible flows: application to heat-driven cavity benchmarks in 2D and 3D (M. BRAACK); comparison of various implicit, explicit, centered and upwind schemes for the simulation of compressed flows on moving mesh (A. CADIOU, M. BUFFAT, L. Le PENVEN, C. Le RIBAULT); low Mach number limit for viscous compressible flows (R. DANCHIN); some Properties of the low Mach number
Chan, David T.; Brauckmann, Gregory J.
2011-01-01
A 6%-scale unpowered model of the Orion Launch Abort Vehicle (LAV) ALAS-11-rev3c configuration was tested in the NASA Langley National Transonic Facility to obtain static aerodynamic data at flight Reynolds numbers. Subsonic and transonic data were obtained for Mach numbers between 0.3 and 0.95 for angles of attack from -4 to +22 degrees and angles of sideslip from -10 to +10 degrees. Data were also obtained at various intermediate Reynolds numbers between 2.5 million and 45 million depending on Mach number in order to examine the effects of Reynolds number on the vehicle. Force and moment data were obtained using a 6-component strain gauge balance that operated both at warm temperatures (+120 . F) and cryogenic temperatures (-250 . F). Surface pressure data were obtained with electronically scanned pressure units housed in heated enclosures designed to survive cryogenic temperatures. Data obtained during the 3-week test entry were used to support development of the LAV aerodynamic database and to support computational fluid dynamics code validation. Furthermore, one of the outcomes of the test was the reduction of database uncertainty on axial force coefficient for the static unpowered LAV. This was accomplished as a result of good data repeatability throughout the test and because of decreased uncertainty on scaling wind tunnel data to flight.
Low Mach Number Fluctuating Hydrodynamics of Multispecies Liquid Mixtures
Donev, A; Bhattacharjee, A K; Garcia, A L; Bell, J B
2014-01-01
We develop a low Mach number formulation of the hydrodynamic equations describing transport of mass and momentum in a multispecies mixture of incompressible miscible liquids at specified temperature and pressure that generalizes our prior work on ideal mixtures of ideal gases and binary liquid mixtures. In this formulation we combine and extend a number of existing descriptions of multispecies transport available in the literature. The formulation applies to non-ideal mixtures of arbitrary number of species, without the need to single out a 'solvent' species, and includes contributions to the diffusive mass flux due to gradients of composition, temperature and pressure. Momentum transport and advective mass transport are handled using a low Mach number approach that eliminates fast sound waves (pressure fluctuations) from the full compressible system of equations and leads to a quasi-incompressible formulation. Thermal fluctuations are included in our fluctuating hydrodynamics description following the princi...
Pendergraft, Odis C., Jr.; Burley, James R., II; Bare, E. Ann
1986-01-01
An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of upper and lower external nozzle flap geometry on the external afterbody/nozzle drag of nonaxisymmetric two-dimensional convergent-divergent exhaust nozzles having parallel external sidewalls installed on a generic twin-engine, fighter-aircraft model. Tests were conducted over a Mach number range from 0.60 to 1.20 and over an angle-of-attack range from -5 to 9 deg. Nozzle pressure ratio was varied from jet off (1.0) to approximately 10.0, depending on Mach number.
Statistical error in particle simulations of low mach number flows
Energy Technology Data Exchange (ETDEWEB)
Hadjiconstantinou, N G; Garcia, A L
2000-11-13
We present predictions for the statistical error due to finite sampling in the presence of thermal fluctuations in molecular simulation algorithms. The expressions are derived using equilibrium statistical mechanics. The results show that the number of samples needed to adequately resolve the flowfield scales as the inverse square of the Mach number. Agreement of the theory with direct Monte Carlo simulations shows that the use of equilibrium theory is justified.
Henneberry, Hugh M.; Snyder, Christopher A.
1993-01-01
An analysis of gas turbine engines using water and oxygen injection to enhance performance by increasing Mach number capability and by increasing thrust is described. The liquids are injected, either separately or together, into the subsonic diffuser ahead of the engine compressor. A turbojet engine and a mixed-flow turbofan engine (MFTF) are examined, and in pursuit of maximum thrust, both engines are fitted with afterburners. The results indicate that water injection alone can extend the performance envelope of both engine types by one and one-half Mach numbers at which point water-air ratios reach 17 or 18 percent and liquid specific impulse is reduced to some 390 to 470 seconds, a level about equal to the impulse of a high energy rocket engine. The envelope can be further extended, but only with increasing sacrifices in liquid specific impulse. Oxygen-airflow ratios as high as 15 percent were investigated for increasing thrust. Using 15 percent oxygen in combination with water injection at high supersonic Mach numbers resulted in thrust augmentation as high as 76 percent without any significant decrease in liquid specific impulse. The stoichiometric afterburner exit temperature increased with increasing oxygen flow, reaching 4822 deg R in the turbojet engine at a Mach number of 3.5. At the transonic Mach number of 0.95 where no water injection is needed, an oxygen-air ratio of 15 percent increased thrust by some 55 percent in both engines, along with a decrease in liquid specific impulse of 62 percent. Afterburner temperature was approximately 4700 deg R at this high thrust condition. Water and/or oxygen injection are simple and straightforward strategies to improve engine performance and they will add little to engine weight. However, if large Mach number and thrust increases are required, liquid flows become significant, so that operation at these conditions will necessarily be of short duration.
Low Mach Number Fluctuating Hydrodynamics of Binary Liquid Mixtures
Nonaka, A J; Bell, J B; Donev, A
2014-01-01
Continuing on our previous work [ArXiv:1212.2644], we develop semi-implicit numerical methods for solving low Mach number fluctuating hydrodynamic equations appropriate for modeling diffusive mixing in isothermal mixtures of fluids with different densities and transport coefficients. We treat viscous dissipation implicitly using a recently-developed variable-coefficient Stokes solver [ArXiv:1308.4605]. This allows us to increase the time step size significantly compared to the earlier explicit temporal integrator. For viscous-dominated flows, such as flows at small scales, we develop a scheme for integrating the overdamped limit of the low Mach equations, in which inertia vanishes and the fluid motion can be described by a steady Stokes equation. We also describe how to incorporate advanced higher-order Godunov advection schemes in the numerical method, allowing for the treatment of fluids with high Schmidt number including the vanishing mass diffusion coefficient limit. We incorporate thermal fluctuations in...
Courant Number and Mach Number Insensitive CE/SE Euler Solvers
Chang, Sin-Chung
2005-01-01
It has been known that the space-time CE/SE method can be used to obtain ID, 2D, and 3D steady and unsteady flow solutions with Mach numbers ranging from 0.0028 to 10. However, it is also known that a CE/SE solution may become overly dissipative when the Mach number is very small. As an initial attempt to remedy this weakness, new 1D Courant number and Mach number insensitive CE/SE Euler solvers are developed using several key concepts underlying the recent successful development of Courant number insensitive CE/SE schemes. Numerical results indicate that the new solvers are capable of resolving crisply a contact discontinuity embedded in a flow with the maximum Mach number = 0.01.
Hysteresis phenomenon of hypersonic inlet at high Mach number
Jiao, Xiaoliang; Chang, Juntao; Wang, Zhongqi; Yu, Daren
2016-11-01
When the hypersonic inlet works at a Mach number higher than the design value, the hypersonic inlet is started with a regular reflection of the external compression shock at the cowl, whereas a Mach reflection will result in the shock propagating forwards to cause a shock detachment at the cowl lip, which is called "local unstart of inlet". As there are two operation modes of hypersonic inlet at high Mach number, the mode transition may occur with the operation condition of hypersonic inlet changing. A cowl-angle-variation-induced hysteresis and a downstream-pressure-variation-induced hysteresis in the hypersonic inlet start↔local unstart transition are obtained by viscous numerical simulations in this paper. The interaction of the external compression shock and boundary layer on the cowl plays a key role in the hysteresis phenomenon. Affected by the transition of external compression shock reflection at the cowl and the transition between separated and attached flow on the cowl, a hysteresis exists in the hypersonic inlet start↔local unstart transition. The hysteresis makes the operation of a hypersonic inlet very difficult to control. In order to avoid hysteresis phenomenon and keep the hypersonic inlet operating in a started mode, the control route should never pass through the local unstarted boundary.
Low Mach number fluctuating hydrodynamics of multispecies liquid mixtures
Energy Technology Data Exchange (ETDEWEB)
Donev, Aleksandar, E-mail: donev@courant.nyu.edu; Bhattacharjee, Amit Kumar [Courant Institute of Mathematical Sciences, New York University, New York, New York 10012 (United States); Nonaka, Andy; Bell, John B. [Center for Computational Science and Engineering, Lawrence Berkeley National Laboratory, Berkeley, California 94720 (United States); Garcia, Alejandro L. [Department of Physics and Astronomy, San Jose State University, San Jose, California 95192 (United States)
2015-03-15
We develop a low Mach number formulation of the hydrodynamic equations describing transport of mass and momentum in a multispecies mixture of incompressible miscible liquids at specified temperature and pressure, which generalizes our prior work on ideal mixtures of ideal gases [Balakrishnan et al., “Fluctuating hydrodynamics of multispecies nonreactive mixtures,” Phys. Rev. E 89 013017 (2014)] and binary liquid mixtures [Donev et al., “Low mach number fluctuating hydrodynamics of diffusively mixing fluids,” Commun. Appl. Math. Comput. Sci. 9(1), 47-105 (2014)]. In this formulation, we combine and extend a number of existing descriptions of multispecies transport available in the literature. The formulation applies to non-ideal mixtures of arbitrary number of species, without the need to single out a “solvent” species, and includes contributions to the diffusive mass flux due to gradients of composition, temperature, and pressure. Momentum transport and advective mass transport are handled using a low Mach number approach that eliminates fast sound waves (pressure fluctuations) from the full compressible system of equations and leads to a quasi-incompressible formulation. Thermal fluctuations are included in our fluctuating hydrodynamics description following the principles of nonequilibrium thermodynamics. We extend the semi-implicit staggered-grid finite-volume numerical method developed in our prior work on binary liquid mixtures [Nonaka et al., “Low mach number fluctuating hydrodynamics of binary liquid mixtures,” http://arxiv.org/abs/1410.2300 (2015)] and use it to study the development of giant nonequilibrium concentration fluctuations in a ternary mixture subjected to a steady concentration gradient. We also numerically study the development of diffusion-driven gravitational instabilities in a ternary mixture and compare our numerical results to recent experimental measurements [Carballido-Landeira et al., “Mixed-mode instability of a
DSMC Simulation of High Mach Number Taylor-Couette Flow
Pradhan, Sahadev, , Dr.
2017-01-01
The main focus of this work is to characterise the Taylor-Couette flow of an ideal gas between two coaxial cylinders at Mach number Ma = (U_w /√{ kbT_w / m }) in the range 0.01 Boltzmann constant. The cylindrical surfaces are specified as being diffusely reflecting with the thermal accommodation coefficient equal to one. In the present analysis of high Mach number compressible Taylor-Couette flow using DSMC method, wall slip in the temperature and the velocities are found to be significant. Slip occurs because the temperature/velocity of the molecules incident on the wall could be very different from that of the wall, even though the temperature/velocity of the reflected molecules is equal to that of the wall. Due to the high surface speed of the inner cylinder, significant heating of the gas is taking place. The gas temperature increases until the heat transfer to the surface equals the work done in moving the surface. The highest temperature is obtained near the moving surface of the inner cylinder at a radius of about (1.26 r_1).
The Variation of Slat Noise with Mach and Reynolds Numbers
Lockhard, David P.; Choudhari, Meelan M.
2011-01-01
The slat noise from the 30P30N high-lift system has been computed using a computational fluid dynamics code in conjunction with a Ffowcs Williams-Hawkings solver. By varying the Mach number from 0.13 to 0.25, the noise was found to vary roughly with the 5th power of the speed. Slight changes in the behavior with directivity angle could easily account for the different speed dependencies reported in the literature. Varying the Reynolds number from 1.4 to 2.4 million resulted in almost no differences, and primarily served to demonstrate the repeatability of the results. However, changing the underlying hybrid Reynolds-averaged-Navier-Stokes/Large-Eddy-Simulation turbulence model significantly altered the mean flow because of changes in the flap separation. However, the general trends observed in both the acoustics and near-field fluctuations were similar for both models.
Low Mach number theory of freely cooling granular gases
Meerson, Baruch; Vilenkin, Arkady
2007-01-01
We use hydrodynamic equations to investigate the dynamics of a freely cooling dilute granular gas with nearly elastic particle collisions. We assume a narrow channel geometry and focus on the regime where the sound travel time through the system is much shorter than the typical cooling time of the gas. As a result, the pressure rapidly becomes almost homogeneous, while the Mach number is small. Eliminating the sound waves and employing Lagrangian coordinates, we reduce the full hydrodynamics to a single nonlinear/nonlocal equation of a reaction-diffusion type. This equation describes a broad class of flows and, in particular, can follow the development of strongly nonlinear states during clustering instability. Without heat diffusion, the reduced equation is exactly soluble and develops a finite-time density blowup with the same local features as those exhibited by the recently found family of exact solutions of the full set of ideal hydrodynamic equations (Fouxon et al. 2007). The heat diffusion, however, ar...
Turbomachinery for Low-to-High Mach Number Flight
Tan, Choon S.; Shah, Parthiv N.
2004-01-01
The thrust capability of turbojet cycles is reduced at high flight Mach number (3+) by the increase in inlet stagnation temperature. The 'hot section' temperature limit imposed by materials technology sets the maximum heat addition and, hence, sets the maximum flight Mach number of the operating envelope. Compressor pre-cooling, either via a heat exchanger or mass-injection, has been suggested as a means to reduce compressor inlet temperature and increase mass flow capability, thereby increasing thrust. To date, however, no research has looked at compressor cooling (i.e., using a compressor both to perform work on the gas path air and extract heat from it simultaneously). We wish to assess the feasibility of this novel concept for use in low-to-high Mach number flight. The results to-date show that an axial compressor with cooling: (1) relieves choking in rear stages (hence opening up operability), (2) yields higher-pressure ratio and (3) yields higher efficiency for a given corrected speed and mass flow. The performance benefit is driven: (i) at the blade passage level, by a decrease in the total pressure reduction coefficient and an increase in the flow turning; and (ii) by the reduction in temperature that results in less work required for a given pressure ratio. The latter is a thermodynamic effect. As an example, calculations were performed for an eight-stage compressor with an adiabatic design pressure ratio of 5. By defining non-dimensional cooling as the percentage of compressor inlet stagnation enthalpy removed by a heat sink, the model shows that a non-dimensional cooling of percent in each blade row of the first two stages can increase the compressor pressure ratio by as much as 10-20 percent. Maximum corrected mass flow at a given corrected speed may increase by as much as 5 percent. In addition, efficiency may increase by as much as 5 points. A framework for characterizing and generating the performance map for a cooled compressor has been developed
The shock tube as a device for testing transonic airfoils at high Reynolds numbers
Cook, W. J.; Presley, L. L.; Chapman, G. T.
1978-01-01
A performance analysis of gas-driven shock tubes shows that transonic airfoil flows with chord Reynolds numbers in the range of 100 million can be generated behind the primary shock in a large shock tube. A study of flow over simple airfoils has been carried out at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photos and airfoil pressure measurements show that steady transonic flows similar to those observed for the airfoils in wind tunnels can be generated within the available testing time in a shock tube with either properly-contoured test section walls or a properly-designed slotted-wall test section. The study indicates that the shock tube is a useful facility for studying two-dimensional high Reynolds number transonic airfoil flows.
Magnus effects on spinning transonic missiles
Seginer, A.; Rosenwasser, I.
1983-01-01
Magnus forces and moments were measured on a basic-finner model spinning in transonic flow. Spin was induced by canted fins or by full-span or semi-span, outboard and inboard roll controls. Magnus force and moment reversals were caused by Mach number, reduced spin rate, and angle of attack variations. Magnus center of pressure was found to be independent of the angle of attack but varied with the Mach number and model configuration or reduced spin rate.
Edge, cavity and aperture tones at very low Mach numbers
Howe, M. S.
1997-01-01
This paper discusses self-sustaining oscillations of high-Reynolds-number shear layers and jets incident on edges and corners at infinitesimal Mach number. These oscillations are frequently sources of narrow-band sound, and are usually attributed to the formation of discrete vortices whose interactions with the edge or corner produce impulsive pressures that lead to the formation of new vorticity and complete a feedback cycle of operation. Linearized analyses of these interactions are presented in which free shear layers are modelled by vortex sheets. Detailed results are given for shear flows over rectangular wall apertures and shallow cavities, and for the classical jet edge interaction. The operating stages of self-sustained oscillations are identified with poles in the upper half of the complex frequency plane of a certain impulse response function. It is argued that the real parts of these poles determine the Strouhal numbers of the operating stages observed experimentally for the real, nonlinear system. The response function coincides with the Rayleigh conductivity of the ‘window’ spanned by the shear flow for wall apertures and jet edge interactions, and to a frequency dependent drag coefficient for shallow wall cavities. When the interaction occurs in the neighbourhood of an acoustic resonator, exemplified by the flue organ pipe, the poles are augmented by a sequence of poles whose real parts are close to the resonance frequencies of the resonator, and the resonator can ‘speak’ at one of these frequencies (by extracting energy from the mean flow) provided the corresponding pole has positive imaginary part.
Capone, F. J.
1982-01-01
An investigation to determine the aeropropulsive characteristics of nonaxisymmetric nozzles on an F-18 jet effects model was conducted in the Langley 16-foot transonic tunnel and the AEDC 16-foot supersonic wind tunnel. The performance of a two dimensional convergent-divergent nozzle, a single expansion ramp nozzle, and a wedge nozzle was compared with that of the baseline axisymmetric nozzle. Test data were obtained at static conditions and at Mach numbers from 0.60 to 2.20 at an angle of attack of 0 deg. Nozzle pressure ratio was varied from jet-off to about 20.
Performance characteristics of two multiaxis thrust-vectoring nozzles at Mach numbers up to 1.28
Wing, David J.; Capone, Francis J.
1993-01-01
The thrust-vectoring axisymmetric (VA) nozzle and a spherical convergent flap (SCF) thrust-vectoring nozzle were tested along with a baseline nonvectoring axisymmetric (NVA) nozzle in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0 to 1.28 and nozzle pressure ratios from 1 to 8. Test parameters included geometric yaw vector angle and unvectored divergent flap length. No pitch vectoring was studied. Nozzle drag, thrust minus drag, yaw thrust vector angle, discharge coefficient, and static thrust performance were measured and analyzed, as well as external static pressure distributions. The NVA nozzle and the VA nozzle displayed higher static thrust performance than the SCF nozzle throughout the nozzle pressure ratio (NPR) range tested. The NVA nozzle had higher overall thrust minus drag than the other nozzles throughout the NPR and Mach number ranges tested. The SCF nozzle had the lowest jet-on nozzle drag of the three nozzles throughout the test conditions. The SCF nozzle provided yaw thrust angles that were equal to the geometric angle and constant with NPR. The VA nozzle achieved yaw thrust vector angles that were significantly higher than the geometric angle but not constant with NPR. Nozzle drag generally increased with increases in thrust vectoring for all the nozzles tested.
DSMC simulations of leading edge flat-plate boundary layer flows at high Mach number
Pradhan, Sahadev, , Dr.
2017-01-01
The flow over a 2D leading-edge flat plate is studied at Mach number Ma = (Uinf /√{kBTinf / m }) in the range Boltzmann constant. The variation of streamwise velocity, temperature, number-density, and mean free path along the wall normal direction away from the plate surface is studied. The qualitative nature of the streamwise velocity at high Mach number is similar to those in the incompressible limit (parabolic profile). However, there are important differences. The amplitudes of the streamwise velocity increase as the Mach number increases and turned into a more flatter profile near the wall. There is significant velocity and temperature slip at the surface of the plate, and the slip increases as the Mach number is increased. It is interesting to note that for the highest Mach numbers considered here, the streamwise velocity at the wall exceeds the sound speed, and the flow is supersonic throughout the flow domain.
Variation with Mach Number of Static and Total Pressures Through Various Screens
Adler, Alfred A
1946-01-01
Tests were conducted in the Langley 24-inch highspeed tunnel to ascertain the static-pressure and total-pressure losses through screens ranging in mesh from 3 to 12 wires per inch and in wire diameter from 0.023 to 0.041 inch. Data were obtained from a Mach number of approximately 0.20 up to the maximum (choking) Mach number obtainable for each screen. The results of this investigation indicate that the pressure losses increase with increasing Mach number until the choking Mach number, which can be computed, is reached. Since choking imposes a restriction on the mass rate of flow and maximum losses are incurred at this condition, great care must be taken in selecting the screen mesh and wire dimmeter for an installation so that the choking Mach number is
A Device for Measuring Sonic Velocity and Compressor Mach Number
1948-07-01
resonator (the only 4 NACA TN No. 1664 accurate measurement required) is measured, as shomn in figure 1, by means of a mercury manometer . The compressor Mach...tube vs not connected to the ccmpressor inlet until after calibration. The pressure in the device was measured by means of the mercury manometer . Fram
Numerical Simulation of Low Mach Number Fluid - Phenomena.
Reitsma, Scott H.
A method for the numerical simulation of low Mach number (M) fluid-acoustic phenomena is developed. This computational fluid-acoustic (CFA) methodology is based upon a set of conservation equations, termed finite-compressible, derived from the unsteady Navier-Stokes equations. The finite-compressible and more familiar pseudo-compressible equations are compared. The impact of derivation assumptions are examined theoretically and through numerical experimentation. The error associated with these simplifications is shown to be of O(M) and proportional to the amplitude of unsteady phenomena. A computer code for the solution of the finite -compressible equations is developed from an existing pseudo -compressible code. Spatial and temporal discretization issues relevant in the context of near field fluid-acoustic simulations are discussed. The finite volume code employs a MUSCL based third order upwind biased flux difference splitting algorithm for the convective terms. An explicit, three stage, second order Runge-Kutta temporal integration is employed for time accurate simulations while an implicit, approximately factored time quadrature is available for steady state convergence acceleration. The CFA methodology is tested in a series of problems which examine the appropriateness of the governing equations, the exacerbation of spatial truncation errors and the degree of temporal accuracy. Characteristic based boundary conditions employing a spatial formulation are developed. An original non-reflective boundary condition based upon the generalization and extension of existing methods is derived and tested in a series of multi-dimensional problems including those involving viscous shear flows and propagating waves. The final numerical experiment is the simulation of boundary layer receptivity to acoustic disturbances. This represents the first simulation of receptivity at a surface inhomogeneity in which the acoustic phenomena is modeled using physically appropriate
A Reynolds Number Study of Wing Leading-Edge Effects on a Supersonic Transport Model at Mach 0.3
Williams, M. Susan; Owens, Lewis R., Jr.; Chu, Julio
1999-01-01
A representative supersonic transport design was tested in the National Transonic Facility (NTF) in its original configuration with small-radius leading-edge flaps and also with modified large-radius inboard leading-edge flaps. Aerodynamic data were obtained over a range of Reynolds numbers at a Mach number of 0.3 and angles of attack up to 16 deg. Increasing the radius of the inboard leading-edge flap delayed nose-up pitching moment to a higher lift coefficient. Deflecting the large-radius leading-edge flap produced an overall decrease in lift coefficient and delayed nose-up pitching moment to even higher angles of attack as compared with the undeflected large- radius leading-edge flap. At angles of attack corresponding to the maximum untrimmed lift-to-drag ratio, lift and drag coefficients decreased while lift-to-drag ratio increased with increasing Reynolds number. At an angle of attack of 13.5 deg., the pitching-moment coefficient was nearly constant with increasing Reynolds number for both the small-radius leading-edge flap and the deflected large-radius leading-edge flap. However, the pitching moment coefficient increased with increasing Reynolds number for the undeflected large-radius leading-edge flap above a chord Reynolds number of about 35 x 10 (exp 6).
Hot-wire anemometry in transonic flow
Horstman, C. C.; Rose, W. C.
1977-01-01
The use of hot-wire anemometry for obtaining fluctuating data in transonic flows has been evaluated. From hot-wire heat loss correlations based on previous transonic data, the sensitivity coefficients for velocity, density, and total temperature fluctuations have been calculated for a wide range of test conditions and sensor parameters. For sensor Reynolds number greater than 20 and high sensor overheat ratios, the velocity sensitivity remains independent of Mach number and equal to the density sensitivity. These conditions were verified by comparisons of predicted sensitivities with those from recent direct calibrations in transonic flows. Based on these results, techniques are presented to obtain meaningful measurements of fluctuating velocity, density, and Reynolds shear stress using hot-wire and hot-film anemometers. Example of these measurements are presented for two transonic boundary layers.
Hot wire anemometry in transonic flow
Horstman, C. C.; Rose, W. C.
1975-01-01
The use of hot-wire anemometry for obtaining fluctuating data in transonic flows has been evaluated. From hot-wire heat loss correlations based on previous transonic data, the sensitivity coefficients for velocity, density, and total temperature fluctuations have been calculated for a wide range of test conditions and sensor parameters. For sensor Reynolds numbers greater than 20 and high sensor overheat ratios, the velocity sensitivity remains independent of Mach number and equal to the density sensitivity. These conclusions were verified by comparisons of predicted sensitivities with those from recent direct calibrations in transonic flows. Based on these results, techniques are presented to obtain meaningful measurements of fluctuating velocity, density, and Reynolds shear stress using hot-wire and hot-film anemometers. Examples of these measurements are presented for two transonic boundary layers.
Numerical studies of transverse curvature effects on transonic flow stability
Macaraeg, M. G.; Daudpota, Q. I.
1992-01-01
A numerical study of transverse curvature effects on compressible flow temporal stability for transonic to low supersonic Mach numbers is presented for axisymmetric modes. The mean flows studied include a similar boundary-layer profile and a nonsimilar axisymmetric boundary-layer solution. The effect of neglecting curvature in the mean flow produces only small quantitative changes in the disturbance growth rate. For transonic Mach numbers (1-1.4) and aerodynamically relevant Reynolds numbers (5000-10,000 based on displacement thickness), the maximum growth rate is found to increase with curvature - the maximum occurring at a nondimensional radius (based on displacement thickness) between 30 and 100.
Transonic Dynamics Tunnel (TDT)
Federal Laboratory Consortium — The Transonic Dynamics Tunnel (TDT) is a continuous flow wind-tunnel facility capable of speeds up to Mach 1.2 at stagnation pressures up to one atmosphere. The TDT...
Application of shock tubes to transonic airfoil testing at high Reynolds numbers
Cook, W. J.; Chaney, M. J.; Presley, L. L.; Chapman, G. T.
1978-01-01
Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers of the order of 100 million can be produced, with limitations being imposed by the structural integrity of the facility or the model. A study of flow development over a simple circular arc airfoil at zero angle of attack was carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls.
Multiobjective Design Optimization of Supersonic Jet Engine in Different Cruise Mach Numbers
Ogawa, Masamichi; Sato, Tetsuya; Kobayashi, Hiroaki; Taguchi, Hideyuki
The aim of this paper is to apply a multi-objective optimization generic algorithm (MOGA) to the conceptual design of the hypersonic/supersonic vehicles with different cruise Mach number. The pre-cooled turbojet engine is employed as a propulsion system and some engine parameters such as the precooler size, compressor size, compression ratio and fuel type are varied in the analysis. The result shows that the optimum cruise Mach number is about 4 if hydrogen fuel is used. Methane fuel instead of hydrogen reduces the vehicle gross weight by 33% in case of the Mach 2 vehicle.
Turbulent mixing of a slightly supercritical Van der Waals fluid at Low-Mach number
Battista, Francesco; Casciola, Carlo Massimo
2014-01-01
Supercritical fluids near the critical point are characterized by liquid-like densities and gas-like transport properties. These features are purposely exploited in different contexts ranging from natural products extraction/fractionation to aerospace propulsion. Large part of studies concerns this last context, focusing on the dynamics of supercritical fluids at high Mach number where compressibility and thermodynamics strictly interact. Despite the widespread use also at low Mach number, the turbulent mixing properties of slightly supercritical fluids have still not investigated in detail in this regime. This topic is addressed here by dealing with Direct Numerical Simulations (DNS) of a coaxial jet of a slightly supercritical Van der Waals fluid. Since acoustic effects are irrelevant in the Low Mach number conditions found in many industrial applications, the numerical model is based on a suitable low-Mach number expansion of the governing equation. According to experimental observations, the weakly superc...
Le, G.; Russell, C. T.; Gosling, J. T.
1994-12-01
We use International Sun-Earth Explorer (ISEE) magnetic field and plasma data to examine dayside magnetopause crossing under conditions of low Mach number and strongly northward interplanetary magnetic field (IMF). When the solar wind Mach number is low, the IMF stregth and magnetoseath field stregth are large, and we expect the effects of magnetic reconection to be the strongest. When the IMF is strongly northward, we find that the location of the magnetopause boundary layer is very stationary in the space, and we observe many features that are common for both typical and low Mach numbers. However, under low Mach number conditions, we have observed some features that would be expected for cusp reconnection. The boundary layer near the subsolar region contains heated magnetosheath plasma with little hot magnetospheric component that has clearly entered the magnetosphere elsewhere. At least some of the structures present in the boundary layer are impulsive. Inside the boundary layer there is also clear evidence of acceleratedflow from the cusp region for strongly northward IMF at low Mach number. Reconnection beyond the cusp can explain the observed field, plasma, and flow signatures. Therefore at low Mach number, reconection is important in the formation of the boundary layer for northward IMF.
Wall-modeled large-eddy simulation of transonic airfoil buffet at high Reynolds number
Fukushima, Yuma; Kawai, Soshi
2016-11-01
In this study, we conduct the wall-modeled large-eddy simulation (LES) of transonic buffet phenomena over the OAT15A supercritical airfoil at high Reynolds number. The transonic airfoil buffet involves shock-turbulent boundary layer interactions and shock vibration associated with the flow separation downstream of the shock wave. The wall-modeled LES developed by Kawai and Larsson PoF (2012) is tuned on the K supercomputer for high-fidelity simulation. We first show the capability of the present wall-modeled LES on the transonic airfoil buffet phenomena and then investigate the detailed flow physics of unsteadiness of shock waves and separated boundary layer interaction phenomena. We also focus on the sustaining mechanism of the buffet phenomena, including the source of the pressure waves propagated from the trailing edge and the interactions between the shock wave and the generated sound waves. This work was supported in part by MEXT as a social and scientific priority issue to be tackled by using post-K computer. Computer resources of the K computer was provided by the RIKEN Advanced Institute for Computational Science (Project ID: hp150254).
Note: A high Mach number arc-driven shock tube for turbulence studies.
Titus, J B; Alexander, A B; Johnson, J A
2013-04-01
A high Mach arc-driven shock tube has been built at the Center for Plasma Science and Technology of Florida A&M University to study shock waves. A larger apparatus with higher voltage was built to study more stable shock waves and subsequent plasmas. Initial measurements of the apparatus conclude that the desired Mach numbers can be reached using only two-thirds the maximum possible energy that the circuit can provide.
Evaluation of transonic wall interference assessment and correction for semi-span wing data
Garriz, Javier A.; Newman, Perry A.; Vatsa, Veer N.; Haigler, Kara J.; Burdges, Kenneth P.
1990-01-01
A newly developed transonic wall interference assessment and correction (WIAC) code is applied to transonic semispan wing data taken in the Lockheed-Georgia Compressible Flow Wind Tunnel (CFWT), in order to evaluate previous WIAC results and corrections. A current state-of-the-art Navier-Stokes free-air code is used as an independent check. Results are presented for Wing C at the transonic edge of its test data matrix in the CFWT. The results demonstrate the effects of wall porosity, Mach number, and angle-of-attack on the transonic wall interference correction.
The small-scale dynamo: Breaking universality at high Mach numbers
Schleicher, Dominik R G; Federrath, Christoph; Bovino, Stefano; Schmidt, Wolfram
2013-01-01
(Abridged) The small-scale dynamo may play a substantial role in magnetizing the Universe under a large range of conditions, including subsonic turbulence at low Mach numbers, highly supersonic turbulence at high Mach numbers and a large range of magnetic Prandtl numbers Pm, i.e. the ratio of kinetic viscosity to magnetic resistivity. Low Mach numbers may in particular lead to the well-known, incompressible Kolmogorov turbulence, while for high Mach numbers, we are in the highly compressible regime, thus close to Burgers turbulence. In this study, we explore whether in this large range of conditions, a universal behavior can be expected. Our starting point are previous investigations in the kinematic regime. Here, analytic studies based on the Kazantsev model have shown that the behavior of the dynamo depends significantly on Pm and the type of turbulence, and numerical simulations indicate a strong dependence of the growth rate on the Mach number of the flow. Once the magnetic field saturates on the current ...
Effect of Mach number on the efficiency of microwave energy deposition in supersonic flow
Lashkov, V. A.; Karpenko, A. G.; Khoronzhuk, R. S.; Mashek, I. Ch.
2016-05-01
The article is devoted to experimental and numerical studies of the efficiency of microwave energy deposition into a supersonic flow around the blunt cylinder at different Mach numbers. Identical conditions for energy deposition have been kept in the experiments, thus allowing to evaluate the pure effect of varying Mach number on the pressure drop. Euler equations are solved numerically to model the corresponding unsteady flow compressed gas. The results of numerical simulations are compared to the data obtained from the physical experiments. It is shown that the momentum, which the body receives during interaction of the gas domain modified by microwave discharge with a shock layer before the body, increases almost linearly with rising of Mach number and the efficiency of energy deposition also rises.
Energy Technology Data Exchange (ETDEWEB)
Dellacherie, St
2004-07-01
This work deals with the derivation of a diphasic low Mach number model obtained through a Mach number asymptotic expansion applied to the compressible diphasic Navier Stokes system, expansion which filters out the acoustic waves. This approach is inspired from the work of Andrew Majda giving the equations of low Mach number combustion for thin flame and for perfect gases. When the equations of state verify some thermodynamic hypothesis, we show that the low Mach number diphasic system predicts in a good way the dilatation or the compression of a bubble and has equilibrium convergence properties. Then, we propose an entropic and convergent Lagrangian scheme in mono-dimensional geometry when the fluids are perfect gases and we propose a first approach in Eulerian variables where the interface between the two fluids is captured with a level set technique. (author)
Sundkvist, David; Krasnoselskikh, V; Bale, S D; Schwartz, S J; Soucek, J; Mozer, F
2012-01-13
Whistler wave trains are observed in the foot region of high Mach number quasiperpendicular shocks. The waves are oblique with respect to the ambient magnetic field as well as the shock normal. The Poynting flux of the waves is directed upstream in the shock normal frame starting from the ramp of the shock. This suggests that the waves are an integral part of the shock structure with the dispersive shock as the source of the waves. These observations lead to the conclusion that the shock ramp structure of supercritical high Mach number shocks is formed as a balance of dispersion and nonlinearity.
Increased Mach Number Capability for the NASA Glenn 10x10 Supersonic Wind Tunnel
Slater, J. W.; Saunders, J. D.
2015-01-01
Computational simulations and wind tunnel testing were conducted to explore the operation of the Abe Silverstein Supersonic Wind Tunnel at the NASA Glenn Research Center at test section Mach numbers above the current limit of Mach 3.5. An increased Mach number would enhance the capability for testing of supersonic and hypersonic propulsion systems. The focus of the explorations was on understanding the flow within the second throat of the tunnel, which is downstream of the test section and is where the supersonic flow decelerates to subsonic flow. Methods of computational fluid dynamics (CFD) were applied to provide details of the shock boundary layer structure and to estimate losses in total pressure. The CFD simulations indicated that the tunnel could be operated up to Mach 4.0 if the minimum width of the second throat was made smaller than that used for previous operation of the tunnel. Wind tunnel testing was able to confirm such operation of the tunnel at Mach 3.6 and 3.7 before a hydraulic failure caused a stop to the testing. CFD simulations performed after the wind tunnel testing showed good agreement with test data consisting of static pressures along the ceiling of the second throat. The CFD analyses showed increased shockwave boundary layer interactions, which was also observed as increased unsteadiness of dynamic pressures collected in the wind tunnel testing.
Carlson, J. R.; Compton, W. B., III
1984-01-01
A wind tunnel investigation was conducted to determine the aerodynamic interference associated with the installation of a long duct, flow-through nacelle on a straight unswept untapered supercritical wing. Experimental data was obtained for the verification of computational prediction techniques. The model was tested in the 16-Foot Transonic Tunnel at Mach numbers from 0.20 to 0.875 and at angles of attack from about 0 deg to 5 deg. The results of the investigation show that strong viscous and compressibility effects are present at the transonic Mach numbers. Numerical comparisons show that linear theory is adequate for subsonic Mach number flow prediction, but is inadequate for prediction of the extreme flow conditions that exist at the transonic Mach numbers.
Mach number study of supersonic turbulence: The properties of the density field
Konstandin, Lukas; Girichidis, Philipp; Peters, Thomas; Shetty, Rahul; Klessen, Ralf S
2015-01-01
We model driven, compressible, isothermal, turbulence with Mach numbers ranging from the subsonic ($\\mathcal{M} \\approx 0.65$) to the highly supersonic regime ($\\mathcal{M}\\approx 16 $). The forcing scheme consists both solenoidal (transverse) and compressive (longitudinal) modes in equal parts. We find a relation $\\sigma_{s}^2 = \\mathrm{b}\\log{(1+\\mathrm{b}^2\\mathcal{M}^2)}$ between the Mach number and the standard deviation of the logarithmic density with $\\mathrm{b} = 0.457 \\pm 0.007$. The density spectra follow $\\mathcal{D}(k,\\,\\mathcal{M}) \\propto k^{\\zeta(\\mathcal{M})}$ with scaling exponents depending on the Mach number. We find $\\zeta(\\mathcal{M}) = \\alpha \\mathcal{M}^{\\beta}$ with a coefficient $\\alpha$ that varies slightly with resolution, whereas $\\beta$ changes systematically. We extrapolate to the limit of infinite resolution and find $\\alpha = -1.91 \\pm 0.01,\\, \\beta =-0.30\\pm 0.03$. The dependence of the scaling exponent on the Mach number implies a fractal dimension $D=2+0.96 \\mathcal{M}^{-0.3...
Tanaka, Kento; Watanabe, Tomoaki; Nagata, Koji; Sasoh, Akihiro; Sakai, Yasuhiko; Hayase, Toshiyuki; Nagoya Univ Collaboration
2016-11-01
The interaction between homogeneous isotropic turbulence and normal shock wave is investigated by direct numerical simulations (DNSs). In the DNSs, a normal shock wave with a shock Mach number 1.1 passes through homogeneous isotropic turbulence with a low turbulent Mach number and a moderate turbulent Reynolds number. The statistics are calculated conditioned on the distance from the shock wave. The results showed that the shock wave makes length scales related to turbulence small. This effect is significant for the Taylor microscale defined with the velocity derivative orthogonal to the shock wave. The decrease in the Kolmogorov scale is also found. Statistics of velocity derivative are found to be changed by the shock wave propagation. The shock wave causes enstrophy amplification due to the dilatation/vorticity interaction. By this interaction, the vorticity components parallel to the shock wave is more amplified than the normal component. The strain rate is also amplified by the shock wave.
Mach number scaling of helicopter rotor blade/vortex interaction noise
Leighton, Kenneth P.; Harris, Wesley L.
1985-01-01
A parametric study of model helicopter rotor blade slap due to blade vortex interaction (BVI) was conducted in a 5 by 7.5-foot anechoic wind tunnel using model helicopter rotors with two, three, and four blades. The results were compared with a previously developed Mach number scaling theory. Three- and four-bladed rotor configurations were found to show very good agreement with the Mach number to the sixth power law for all conditions tested. A reduction of conditions for which BVI blade slap is detected was observed for three-bladed rotors when compared to the two-bladed baseline. The advance ratio boundaries of the four-bladed rotor exhibited an angular dependence not present for the two-bladed configuration. The upper limits for the advance ratio boundaries of the four-bladed rotors increased with increasing rotational speed.
Two-dimensional lattice Boltzmann model for compressible flows with high Mach number
Gan, Yanbiao; Xu, Aiguo; Zhang, Guangcai; Yu, Xijun; Li, Yingjun
2008-03-01
In this paper we present an improved lattice Boltzmann model for compressible Navier-Stokes system with high Mach number. The model is composed of three components: (i) the discrete-velocity-model by M. Watari and M. Tsutahara [Phys. Rev. E 67 (2003) 036306], (ii) a modified Lax-Wendroff finite difference scheme where reasonable dissipation and dispersion are naturally included, (iii) artificial viscosity. The improved model is convenient to compromise the high accuracy and stability. The included dispersion term can effectively reduce the numerical oscillation at discontinuity. The added artificial viscosity helps the scheme to satisfy the von Neumann stability condition. Shock tubes and shock reflections are used to validate the new scheme. In our numerical tests the Mach numbers are successfully increased up to 20 or higher. The flexibility of the new model makes it suitable for tracking shock waves with high accuracy and for investigating nonlinear nonequilibrium complex systems.
Flow-induced cylinder noise formulated as a diffraction problem for low Mach numbers
Gloerfelt, X.; Pérot, F.; Bailly, C.; Juvé, D.
2005-10-01
The role of surfaces in the mechanism of sound generation by low Mach number flows interacting with solid nonvibrating surfaces is well established by the classical aeroacoustic papers by Powell, Doak, Ffowcs Williams, Crighton, or Howe. It can be formulated as a problem of diffraction of the flow sources by the rigid body. The present study illustrates this statement in the case of flow-induced cylinder noise. Curle's formulation is analytically and numerically compared to a formulation based on an exact Green's function tailored to a cylindrical geometry. The surface integral of Curle's formulation represents exactly the diffraction effects by the rigid body. The direct and scattered parts of the sound field are studied. In this low Mach number configuration, the cylinder is compact, and the scattered (dipole) field dominates the direct (quadrupole) field. The classical properties of the scattering by a cylinder are retrieved by considering a point quadripole source near the cylinder surface.
Extension of the pressure correction method to zero-Mach number compressible flows
Institute of Scientific and Technical Information of China (English)
无
2009-01-01
In the present paper,the classical pressure correction method was extended into low Mach number compressible flow regime by integrating equation of state into SIMPLE algorithm.The self-developed code based on this algorithm was applied to predicting the lid-driven cavity flow and shock tube prob-lems,and the results showed good agreement with benchmark solutions and the Mach number can reach the magnitude of as low as 10-5.The attenuation of sound waves in viscous medium was then simulated.The results agree well with the analytical solutions given by theoretical acoustics.This demonstrated that the present method could also be implemented in acoustics field simulation,which is crucial for thermoacoustic simulation.
The Density Variance--Mach Number Relation in Supersonic Turbulence: I. Isothermal, magnetised gas
Molina, F Z; Federrath, C; Klessen, R S
2012-01-01
It is widely accepted that supersonic, magnetised turbulence plays a fundamental role for star formation in molecular clouds. It produces the initial dense gas seeds out of which new stars can form. However, the exact relation between gas compression, turbulent Mach number, and magnetic field strength is still poorly understood. Here, we introduce and test an analytical prediction for the relation between the density variance and the root-mean-square Mach number in supersonic, isothermal, magnetised turbulent flows. We approximate the density and velocity structure of the interstellar medium as a superposition of shock waves. We obtain the density contrast considering the momentum continuity equation for a single magnetised shock and extrapolate this result to the entire cloud. Depending on the field geometry, we then make three different assumptions based on observational and theoretical constraints: B independent of density, B proportional to the root square of the density and B proportional to the density....
Extension of the pressure correction method to zero-Mach number compressible flows
Institute of Scientific and Technical Information of China (English)
HE YaLing; HUANG Jing; TAO YuBing; TAO WenQuan
2009-01-01
In the present paper, the classical pressure correction method was extended into low Mach number compressible flow regime by integrating equation of state into SIMPLE algorithm. The self-developed code based on this algorithm was applied to predicting the lid-driven cavity flow and shock tube prob-lems, and the results showed good agreement with benchmark solutions and the Mach number can reach the magnitude of as low as 10-5. The attenuation of sound waves in viscous medium was then simulated. The results agree well with the analytical solutions given by theoretical acoustics. This demonstrated that the present method could also be implemented in acoustics field simulation, which is crucial for thermoacoustic simulation.
Axisymmetric vortex method for low-Mach number, diffusion-controlled combustion
Lakkis, I
2003-01-01
A grid-free, Lagrangian method for the accurate simulation of low-Mach number, variable-density, diffusion-controlled reacting flow is presented. A fast-chemistry model in which the conversion rate of reactants to products is limited by the local mixing rate is assumed in order to reduce the combustion problem to the solution of a convection-diffusion-generation equation with volumetric expansion and vorticity generation at the reaction fronts. The solutions of the continuity and vorticity equations, and the equations governing the transport of species and energy, are obtained using a formulation in which particles transport conserved quantities by convection and diffusion. The dynamic impact of exothermic combustion is captured through accurate integration of source terms in the vorticity transport equations at the location of the particles, and the extra velocity field associated with volumetric expansion at low Mach number computed to enforced mass conservation. The formulation is obtained for an axisymmet...
Low Mach and Peclet number limit for a model of stellar tachocline and upper radiative zones
Directory of Open Access Journals (Sweden)
Donatella Donatelli
2016-09-01
Full Text Available We study a hydrodynamical model describing the motion of internal stellar layers based on compressible Navier-Stokes-Fourier-Poisson system. We suppose that the medium is electrically charged, we include energy exchanges through radiative transfer and we assume that the system is rotating. We analyze the singular limit of this system when the Mach number, the Alfven number, the Peclet number and the Froude number approache zero in a certain way and prove convergence to a 3D incompressible MHD system with a stationary linear transport equation for transport of radiation intensity. Finally, we show that the energy equation reduces to a steady equation for the temperature corrector.
Nearfield Unsteady Pressures at Cruise Mach Numbers for a Model Scale Counter-Rotation Open Rotor
Stephens, David B.
2012-01-01
An open rotor experiment was conducted at cruise Mach numbers and the unsteady pressure in the nearfield was measured. The system included extensive performance measurements, which can help provide insight into the noise generating mechanisms in the absence of flow measurements. A set of data acquired at a constant blade pitch angle but various rotor speeds was examined. The tone levels generated by the front and rear rotor were found to be nearly equal when the thrust was evenly balanced between rotors.
Bispen, Georgij; Lukáčová-Medvid'ová, Mária; Yelash, Leonid
2017-04-01
In this paper we will present and analyze a new class of the IMEX finite volume schemes for the Euler equations with a gravity source term. We will in particular concentrate on a singular limit of weakly compressible flows when the Mach number M ≪ 1. In order to efficiently resolve slow dynamics we split the whole nonlinear system in a stiff linear part governing the acoustic and gravity waves and a non-stiff nonlinear part that models nonlinear advection effects. For time discretization we use a special class of the so-called globally stiffly accurate IMEX schemes and approximate the stiff linear operator implicitly and the non-stiff nonlinear operator explicitly. For spatial discretization the finite volume approximation is used with the central and Rusanov/Lax-Friedrichs numerical fluxes for the linear and nonlinear subsystem, respectively. In the case of a constant background potential temperature we prove theoretically that the method is asymptotically consistent and asymptotically stable uniformly with respect to small Mach number. We also analyze experimentally convergence rates in the singular limit when the Mach number tends to zero.
Particle-in-cell simulations of particle energization from low Mach number fast mode shocks
Park, Jaehong; Blackman, Eric G; Ren, Chuang; Siller, Robert
2012-01-01
Astrophysical shocks are often studied in the high Mach number limit but weakly compressive fast shocks can occur in magnetic reconnection outflows and are considered to be a site of particle energization in solar flares. Here we study the microphysics of such perpendicular, low Mach number collisionless shocks using two-dimensional particle-in-cell (PIC) simulations with a reduced ion/electron mass ratio and employ a moving wall boundary method for initial generation of the shock. This moving wall method allows for more control of the shock speed, smaller simulation box sizes, and longer simulation times than the commonly used fixed wall, reflection method of shock formation. Our results, which are independent of the shock formation method, reveal the prevalence shock drift acceleration (SDA) of both electron and ions in a purely perpendicular shock with Alfv\\'en Mach number $M_A=6.8$ and ratio of thermal to magnetic pressure $\\beta=8$. We determine the respective minimum energies required for electrons and ...
A half-explicit, non-split projection method for low Mach number flows.
Energy Technology Data Exchange (ETDEWEB)
Pousin, Jerome G. (National Institute for Applied Sciences, France); Najm, Habib N.; Pebay, Philippe Pierre
2004-02-01
In the context of the direct numerical simulation of low MACH number reacting flows, the aim of this article is to propose a new approach based on the integration of the original differential algebraic (DAE) system of governing equations, without further differentiation. In order to do so, while preserving a possibility of easy parallelization, it is proposed to use a one-step index 2 DAE time-integrator, the Half Explicit Method (HEM). In this context, we recall why the low MACH number approximation belongs to the class of index 2 DAEs and discuss why the pressure can be associated with the constraint. We then focus on a fourth-order HEM scheme, and provide a formulation that makes its implementation more convenient. Practical details about the consistency of initial conditions are discussed, prior to focusing on the implicit solve involved in the method. The method is then evaluated using the Modified KAPS Problem, since it has some of the features of the low MACH number approximation. Numerical results are presented, confirming the above expectations. A brief summary of ongoing efforts is finally provided.
The Dynamics of Very High Alfvén Mach Number Shocks in Space Plasmas
Sundberg, Torbjörn; Burgess, David; Scholer, Manfred; Masters, Adam; Sulaiman, Ali H.
2017-02-01
Astrophysical shocks, such as planetary bow shocks or supernova remnant shocks, are often in the high or very-high Mach number regime, and the structure of such shocks is crucial for understanding particle acceleration and plasma heating, as well inherently interesting. Recent magnetic field observations at Saturn’s bow shock, for Alfvén Mach numbers greater than about 25, have provided evidence for periodic non-stationarity, although the details of the ion- and electron-scale processes remain unclear due to limited plasma data. High-resolution, multi-spacecraft data are available for the terrestrial bow shock, but here the very high Mach number regime is only attained on extremely rare occasions. Here we present magnetic field and particle data from three such quasi-perpendicular shock crossings observed by the four-spacecraft Cluster mission. Although both ion reflection and the shock profile are modulated at the upstream ion gyroperiod timescale, the dominant wave growth in the foot takes place at sub-proton length scales and is consistent with being driven by the ion Weibel instability. The observed large-scale behavior depends strongly on cross-scale coupling between ion and electron processes, with ion reflection never fully suppressed, and this suggests a model of the shock dynamics that is in conflict with previous models of non-stationarity. Thus, the observations offer insight into the conditions prevalent in many inaccessible astrophysical environments, and provide important constraints for acceleration processes at such shocks.
Agarwal, Shankar
2013-01-01
We calculate the cosmic Mach number M - the ratio of the bulk flow of the velocity field on scale R to the velocity dispersion within regions of scale R. M is effectively a measure of the ratio of large-scale to small-scale power and can be a useful tool to constrain the cosmological parameter space. Using a compilation of existing peculiar velocity surveys, we calculate M and compare it to that estimated from mock catalogues extracted from the LasDamas (a LCDM cosmology) numerical simulations. We find agreement with expectations for the LasDamas cosmology at ~ 1.5 sigma CL. We also show that our Mach estimates for the mocks are not biased by selection function effects. To achieve this, we extract dense and nearly-isotropic distributions using Gaussian selection functions with the same width as the characteristic depth of the real surveys, and show that the Mach numbers estimated from the mocks are very similar to the values based on Gaussian profiles of the corresponding widths. We discuss the importance of ...
Bonhaus, Daryl L.; Maddalon, Dal V.
1998-01-01
Flight-measured high Reynolds number turbulent-flow pressure distributions on a transport wing in transonic flow are compared to unstructured-grid calculations to assess the predictive ability of a three-dimensional Euler code (USM3D) coupled to an interacting boundary layer module. The two experimental pressure distributions selected for comparative analysis with the calculations are complex and turbulent but typical of an advanced technology laminar flow wing. An advancing front method (VGRID) was used to generate several tetrahedral grids for each test case. Initial calculations left considerable room for improvement in accuracy. Studies were then made of experimental errors, transition location, viscous effects, nacelle flow modeling, number and placement of spanwise boundary layer stations, and grid resolution. The most significant improvements in the accuracy of the calculations were gained by improvement of the nacelle flow model and by refinement of the computational grid. Final calculations yield results in close agreement with the experiment. Indications are that further grid refinement would produce additional improvement but would require more computer memory than is available. The appendix data compare the experimental attachment line location with calculations for different grid sizes. Good agreement is obtained between the experimental and calculated attachment line locations.
Dixon, G. V.; Barringer, S. R.; Gray, C. E.; Leatherman, A. D.
1975-01-01
Computer programs and resulting tabulations are presented of pipeline length-to-diameter ratios as a function of Mach number and pressure ratios for compressible flow. The tabulations are applicable to air, nitrogen, oxygen, and hydrogen for compressible isothermal flow with friction and compressible adiabatic flow with friction. Also included are equations for the determination of weight flow. The tabulations presented cover a wider range of Mach numbers for choked, adiabatic flow than available from commonly used engineering literature. Additional information presented, but which is not available from this literature, is unchoked, adiabatic flow over a wide range of Mach numbers, and choked and unchoked, isothermal flow for a wide range of Mach numbers.
Triple-deck analysis of transonic high Reynolds number flow through slender channels.
Kluwick, A; Kornfeld, M
2014-07-28
In this work, laminar transonic weakly three-dimensional flows at high Reynolds numbers in slender channels, as found in microsupersonic nozzles and turbomachines of micro-electro-mechanical systems, are considered. The channel height is taken so small that the viscous wall layers forming at the channel walls start to interact strongly rather than weakly with the inviscid core flow and, therefore, the classical boundary layer approach fails. The resulting viscous-inviscid interaction problem is formulated using matched asymptotic expansions and found to be governed by a triple-deck structure. As a consequence, the properties of the predominantly inviscid core region and the viscous wall layers have to be calculated simultaneously in the interaction region. Weakly three-dimensional effects caused by surface roughness, upstream propagating flow perturbations, boundary layer separation as well as bifurcating solutions are discussed. Representative results for subsonic as well as supersonic conditions are presented, and the importance of these flow phenomena in technical applications as, for example, a means to reduce shock losses through the use of deformed geometry is addressed.
Federrath, Christoph; Schober, Jennifer; Banerjee, Robi; Klessen, Ralf S; Schleicher, Dominik R G; 10.1103/PhysRevLett.107.114504
2011-01-01
We study the growth rate and saturation level of the turbulent dynamo in magnetohydrodynamical simulations of turbulence, driven with solenoidal (divergence-free) or compressive (curl-free) forcing. For models with Mach numbers ranging from 0.02 to 20, we find significantly different magnetic field geometries, amplification rates, and saturation levels, decreasing strongly at the transition from subsonic to supersonic flows, due to the development of shocks. Both extreme types of turbulent forcing drive the dynamo, but solenoidal forcing is more efficient, because it produces more vorticity.
The Experimental Measurement of Aerodynamic Heating About Complex Shapes at Supersonic Mach Numbers
Neumann, Richard D.; Freeman, Delma C.
2011-01-01
In 2008 a wind tunnel test program was implemented to update the experimental data available for predicting protuberance heating at supersonic Mach numbers. For this test the Langley Unitary Wind Tunnel was also used. The significant differences for this current test were the advances in the state-of-the-art in model design, fabrication techniques, instrumentation and data acquisition capabilities. This current paper provides a focused discussion of the results of an in depth analysis of unique measurements of recovery temperature obtained during the test.
Convective heat transport in stratified atmospheres at low and high Mach number
Anders, Evan H
2016-01-01
Convection in astrophysical systems is stratified and often occurs at high Rayleigh number (Ra) and low Mach number (Ma). Here we study stratified convection in the context of plane-parallel, polytropically stratified atmospheres. We hold the density stratification ($n_{\\rho}$) and Prandtl number (Pr) constant while varying Ma and Ra to determine the behavior of the Nusselt number (Nu), which quantifies the efficiency of convective heat transport. As Ra increases and $\\text{Ma} \\rightarrow 1$, a scaling of Nu $\\propto$ Ra$^{0.45}$ is observed. As Ra increases to a regime where Ma $\\geq 1$, this scaling gives way to a weaker Nu $\\propto$ Ra$^{0.19}$. In the regime of Ma $\\ll 1$, a consistent Nu $\\propto$ Ra$^{0.31}$ is retrieved, reminiscent of the Nu $\\propto$ Ra$^{2/7}$ seen in Rayleigh-B\\'{e}nard convection.
Effects of nonuniform Mach-number entrance on scramjet nozzle flowfield and performance
Zhang, Pu; Xu, Jinglei; Quan, Zhibin; Mo, Jianwei
2016-12-01
Considering the non-uniformities of nozzle entrance influenced by the upstream, the effects of nonuniform Mach-number coupled with shock and expansion-wave on the flowfield and performances of single expansion ramp nozzle (SERN) are numerically studied using Reynolds-Averaged Navier-Stokes equations. The adopted Reynolds-averaged Navier-Stokes methodology is validated by comparing the numerical results with the cold experimental data, and the average method used in this paper is discussed. Uniform and nonuniform facility nozzles are designed to generate different Mach-number profile for the inlet of SERN, which is direct-connected with different facility nozzle, and the whole flowfield is simulated. Because of the coupling of shock and expansion-wave, flow direction of nonuniform SERN entrance is distorted. Compared with Mach contour of uniform case, the line is more curved for coupling shock-wave entrance (SWE) case, and flatter for the coupling expansion-wave entrance (EWE) case. Wall pressure distribution of SWE case appears rising region, whereas decreases like stairs of EWE case. The numerical results reveal that the coupled shock and expansion-wave play significant roles on nozzle performances. Compared with the SERN performances of uniform entrance case at the same work conditions, the thrust of nonuniform entrance cases reduces by 3-6%, pitch moment decreases by 2.5-7%. The negative lift presents an incremental trend with EWE while the situation is the opposite with SWE. These results confirm that considering the entrance flow parameter nonuniformities of a scramjet nozzle coupled with shock or expansion-wave from the upstream is necessary.
Unsteady transonic flow in cascades
Surampudi, S. P.; Adamczyk, J. J.
1984-01-01
There is a need for methods to predict the unsteady air loads associated with flutter of turbomachinery blading at transonic speeds. The results of such an analysis in which the steady relative flow approaching a cascade of thin airfoils is assumed to be transonic, irrotational, and isentropic is presented. The blades in the cascade are allowed to undergo a small amplitude harmonic oscillation which generates a small unsteady flow superimposed on the existing steady flow. The blades are assumed to oscillate with a prescribed motion of constant amplitude and interblade phase angle. The equations of motion are obtained by linearizing about a uniform flow the inviscid nonheat conducting continuity and momentum equations. The resulting equations are solved by employing the Weiner Hopf technique. The solution yields the unsteady aerodynamic forces acting on the cascade at Mach number equal to 1. Making use of an unsteady transonic similarity law, these results are compared with the results obtained from linear unsteady subsonic and supersonic cascade theories. A parametric study is conducted to find the effects of reduced frequency, solidity, stagger angle, and position of pitching axis on the flutter.
基于预处理HLLEW格式的全速域数值算法%Preconditioning HLLEW Scheme for Flows at All Mach Numbers
Institute of Scientific and Technical Information of China (English)
刘中玉; 张明锋; 郑冠男; 杨国伟
2016-01-01
Based on HLLEW ( Harten⁃Lax⁃Van Leer⁃Einfeldt⁃Wada) scheme, low speed preconditioning technology is introduced to develop a three⁃dimensional Navier⁃Stokes solver for flows at all Mach numbers. Low speed preconditioning techniques is introduced to reconstruct dissipative term in HLLEW scheme and preconditioning HLLEW scheme is proposed. Implicit time⁃marching method is constructed based on preconditioning Jacobian Matrix. Results of NACA 4412 incompressible flow and RAE 2822 transonic flow with preconditioning HLLEW scheme are compared with results by original method and experimental data. It shows that preconditioning HLLEW method improves accuracy and convergence rate for low speed flow. It can be applied for flows at all Mach numbers.%基于HLLEW（ Harten⁃Lax⁃Van Leer⁃Einfeldt⁃Wada）格式引入预处理技术发展适合求解全速域流场的三维Navier⁃Stokes求解器。引入低速预处理技术，重新构造HLLEW格式的耗散项，给出预处理后的HLLEW格式，并根据预处理后的雅克比矩阵构造相应的隐式时间推进方程。利用预处理方法求解 NACA 4412低速不可压流动与RAE 2822跨声速可压缩流动，并与实验结果及原有方法的计算结果对比。结果表明：预处理HLLEW格式不仅提高低速不可压缩流动的计算效率和精度，也保持了对可压缩流动的处理能力，是一种适用于全速域流场数值模拟的有效方法。
Specularly reflected He sup 2+ at high Mach number quasi-parallel shocks
Energy Technology Data Exchange (ETDEWEB)
Fuselier, S.A.; Lennartsson, O.W. (Lockheed Palo Alto Research Lab., CA (United States)); Thomsen, M.F. (Los Alamos National Lab., NM (United States)); Russell, C.T. (Univ. of California, Los Angeles (United States))
1990-04-01
Upstream from the Earth's quasi-parallel bow shock, the Lockheed Plasma Composition Experiment on ISEE 1 often observes two types of suprathermal He{sup 2+} distributions. Always present to some degree is an energetic (several keV/eto 17.4 keV/e, the maximum energy of the detector) diffuse He{sup 2+} distribution. Sometimes, apparently when the Alfven Mach number, M{sub A}, is high enough and the spacecraft is near the shock (within a few minutes of a crossing), a second type of suprathermal He{sup 2+} distribution is also observed. This nongyrotropic, gyrating He{sup 2+} distribution has velocity components parallel and perpendicular to the magnetic field that are consistent with near-specular reflection of a portion of the incident solar wind He{sup 2+} distribution off the shock. Specularly reflected and diffuse proton distributions are associated with these gyrating He{sup 2+} distributions. The presence of these gyrating He{sup 2+} distributions suggests that specular reflection is controlled primarily by magnetic forces in high Mach number quasi-parallel shocks and that these distributions may be a seed population for more energetic diffuse He{sup 2+} distributions.
The influence of incident shock Mach number on radial incident shock wave focusing
Directory of Open Access Journals (Sweden)
Xin Chen
2016-04-01
Full Text Available Experiments and numerical simulations were carried out to investigate radial incident shock focusing on a test section where the planar incident shock wave was divided into two identical ones. A conventional shock tube was used to generate the planar shock. Incident shock Mach number of 1.51, 1.84 and 2.18 were tested. CCD camera was used to obtain the schlieren photos of the flow field. Third-order, three step strong-stability-preserving (SSP Runge-Kutta method, third-order weighed essential non-oscillation (WENO scheme and adaptive mesh refinement (AMR algorithm were adopted to simulate the complicated flow fields characterized by shock wave interaction. Good agreement between experimental and numerical results was observed. Complex shock wave configurations and interactions (such as shock reflection, shock-vortex interaction and shock focusing were observed in both the experiments and numerical results. Some new features were observed and discussed. The differences of structure of flow field and the variation trends of pressure were compared and analyzed under the condition of different Mach numbers while shock wave focusing.
The density variance - Mach number relation in isothermal and non-isothermal adiabatic turbulence
Nolan, Chris A; Sutherland, Ralph S
2015-01-01
The density variance - Mach number relation of the turbulent interstellar medium is relevant for theoretical models of the star formation rate, efficiency, and the initial mass function of stars. Here we use high-resolution hydrodynamical simulations with grid resolutions of up to 1024^3 cells to model compressible turbulence in a regime similar to the observed interstellar medium. We use Fyris Alpha, a shock-capturing code employing a high-order Godunov scheme to track large density variations induced by shocks. We investigate the robustness of the standard relation between the logarithmic density variance (sigma_s^2) and the sonic Mach number (M) of isothermal interstellar turbulence, in the non-isothermal regime. Specifically, we test ideal gases with diatomic molecular (gamma = 7/5) and monatomic (gamma = 5/3) adiabatic indices. A periodic cube of gas is stirred with purely solenoidal forcing at low wavenumbers, leading to a fully-developed turbulent medium. We find that as the gas heats in adiabatic comp...
A NOVEL SLIGHTLY COMPRESSIBLE MODEL FOR LOW MACH NUMBER PERFECT GAS FLOW CALCULATION
Institute of Scientific and Technical Information of China (English)
邓小刚; 庄逢甘
2002-01-01
By analyzing the characteristics of low Mach number perfect gas flows, a novel Slightly Compressible Model (SCM) for low Mach number perfect gas flows is derived. In view of numerical calculations, this model is proved very efficient,for it is kept within the p-v frame but does not have to satisfy the time consuming divergence-free condition in order to get the incompressible Navier-Stokes equation solutions. Writing the equations in the form of conservation laws, we have derived the characteristic systems which are necessary for numerical calculations. A cellcentered finite-volume method with flux difference upwind-biased schemes is used for the equation solutions and a new Exact Newton Relaxation (ENR) implicit method is developed. Various computed results are presented to validate the present model.Laminar flow solutions over a circular cylinder with wake developing and vortex shedding are presented. Results for inviscid flow over a sphere are compared in excellent agreement with the exact analytic incompressible solution. Three-dimensional viscous flow solutions over sphere and prolate spheroid are also calculated and compared well with experiments and other incompressible solutions. Finally, good convergent performaces are shown for sphere viscous flows.
Airfoil Aeroelastic Flutter Analysis Based on Modified Leishman-Beddoes Model at Low Mach Number
Institute of Scientific and Technical Information of China (English)
SHAO Song; ZHU Qinghua; ZHANG Chenglin; NI Xianping
2011-01-01
Based on modified Leishman-Beddoes(L-B)state space model at low Mach number(lower than 0.3),the airfoil aeroelastic system is presented in this paper.The main modifications for L-B model include a new dynamic stall criterion and revisions of normal force and pitching moment coefficient.The bifurcation diagrams,the limit cycle oscillation (LCO)phase plane plots and the time domain response figures are applied to investigating the stall flutter bifurcation behavior of airfoil aeroelastic systems with symmetry or asymmetry.It is shown that the symmetric periodical oscillation happens after subcritical bifurcation caused by dynamic stall,and the asymmetric periodical oscillation,which is caused by the interaction of dynamic stall and static divergence,only happens in the airfoil aeroelastic system with asymmetry.Validations of the modified L-B model and the airfoil aeroelastic system are presented with the experimental airload data of NACA0012 and OA207 and experimental stall flutter data of NACA0012 respectively.Results demonstrate that the airfoil aeroelastic system presented in this paper is effective and accurate,which can be applied to the investigation of airfoil stall flutter at low Mach number.
Determination of instantaneous pressure in a transonic base flow using four-pulse tomographic PIV
Blinde, P.L.; Lynch, K.P.; Schrijer, F.F.J.; Van Oudheusden, B.W.
2015-01-01
A tomographic four-pulse PIV system is used in a transonic axisymmetric base flow experiment at a nominal free stream Mach number of 0.7, with the objective to obtain flow acceleration and pressure data. The PIV system, consisting of two double-pulse lasers and twelve cameras, allows acquiring two v
Opacity Broadening of $^{13}$CO Linewidths and its Effect on the Variance-Sonic Mach Number Relation
Correia, Caio; Lazarian, Alex; Ossenkopf, Volker; Stutzki, Jürgen; Kainulainen, Jouni; Kowal, Grzegorz; de Medeiros, José Renan
2014-01-01
We study how the estimation of the sonic Mach number ($M_s$) from $^{13}$CO linewidths relates to the actual 3D sonic Mach number. For this purpose we analyze MHD simulations which include post-processing to take radiative transfer effects into account. As expected, we find very good agreement between the linewidth estimated sonic Mach number and the actual sonic Mach number of the simulations for optically thin tracers. However, we find that opacity broadening causes $M_s$ to be overestimated by a factor of ~ 1.16 - 1.3 when calculated from optically thick $^{13}$CO lines. We also find that there is a dependency on the magnetic field: super-Alfv\\'enic turbulence shows increased line broadening as compared with sub-Alfv\\'enic turbulence for all values of optical depth for supersonic turbulence. Our results have implications for the observationally derived sonic Mach number--density standard deviation ($\\sigma_{\\rho/}$) relationship, $\\sigma^2_{\\rho/}=b^2M_s^2$, and the related column density standard deviatio...
Tests of Full-Scale Helicopter Rotors at High Advancing Tip Mach Numbers and Advance Ratios
Biggers, James C.; McCloud, John L., III; Stroub, Robert H.
2015-01-01
As a continuation of the studies of reference 1, three full-scale helicopter rotors have been tested in the Ames Research Center 40- by SO-foot wind tunnel. All three of them were two-bladed, teetering rotors. One of the rotors incorporated the NACA 0012 airfoil section over the entire length of the blade. This rotor was tested at advance ratios up to 1.05. Both of the other rotors were tapered in thickness and incorporated leading-edge camber over the outer 20 percent of the blade radius. The larger of these rotors was tested at advancing tip Mach numbers up to 1.02. Data were obtained for a wide range of lift and propulsive force, and are presented without discussion.
On the proper Mach number and ratio of specific heats for modeling the Venus bow shock
Tatrallyay, M.; Russell, C. T.; Luhmann, J. G.; Barnes, A.; Mihalov, J. D.
1984-01-01
Observational data from the Pioneer Venus Orbiter are used to investigate the physical characteristics of the Venus bow shock, and to explore some general issues in the numerical simulation of collisionless shocks. It is found that since equations from gas-dynamic (GD) models of the Venus shock cannot in general replace MHD equations, it is not immediately obvious what the optimum way is to describe the desired MHD situation with a GD code. Test case analysis shows that for quasi-perpendicular shocks it is safest to use the magnetospheric Mach number as an input to the GD code. It is also shown that when comparing GD predicted temperatures with MHD predicted temperatures total energy should be compared since the magnetic energy density provides a significant fraction of the internal energy of the MHD fluid for typical solar wind parameters. Some conclusions are also offered on the properties of the terrestrial shock.
Relativistic Electron Shock Drift Acceleration in Low Mach Number Galaxy Cluster Shocks
Matsukiyo, Shuichi; Yamazaki, Ryo; Umeda, Takayuki
2011-01-01
An extreme case of electron shock drift acceleration in low Mach number collisionless shocks is investigated as a plausible mechanism of initial acceleration of relativistic electrons in large-scale shocks in galaxy clusters where upstream plasma temperature is of the order of 10 keV and a degree of magnetization is not too small. One-dimensional electromagnetic full particle simulations reveal that, even though a shock is rather moderate, a part of thermal incoming electrons are accelerated and reflected through relativistic shock drift acceleration and form a local nonthermal population just upstream of the shock. The accelerated electrons can self-generate local coherent waves and further be back-scattered toward the shock by those waves. This may be a scenario for the first stage of the electron shock acceleration occurring at the large-scale shocks in galaxy clusters such as CIZA J2242.8+5301 which has well defined radio relics.
Low-Mach-number turbulence in interstellar gas revealed by radio polarization gradients
Gaensler, Bryan M; Burkhart, Blakesley; Newton-McGee, Katherine J; Ekers, Ronald D; Lazarian, Alex; McClure-Griffiths, Naomi M; Robishaw, Timothy; Dickey, John M; Green, Anne J; 10.1038/nature10446
2011-01-01
The interstellar medium of the Milky Way is multi-phase, magnetized and turbulent. Turbulence in the interstellar medium produces a global cascade of random gas motions, spanning scales ranging from 100 parsecs to 1000 kilometres. Fundamental parameters of interstellar turbulence such as the sonic Mach number (the speed of sound) have been difficult to determine because observations have lacked the sensitivity and resolution to directly image the small-scale structure associated with turbulent motion. Observations of linear polarization and Faraday rotation in radio emission from the Milky Way have identified unusual polarized structures that often have no counterparts in the total radiation intensity or at other wavelengths, and whose physical significance has been unclear. Here we report that the gradient of the Stokes vector (Q,U), where Q and U are parameters describing the polarization state of radiation, provides an image of magnetized turbulence in diffuse ionized gas, manifested as a complex filamenta...
Varsakelis, Christos; Papalexandris, Miltiadis V.
2017-01-01
A conundrum in non-equilibrium thermodynamics of heterogeneous mixtures with microstructure concerns the selection of thermodynamic currents and forces in the entropy production rate from the multitude of available options. The objective of this article is to demonstrate that the low-Mach-number approximation can narrow down this ambiguity. More specifically, by postulating that the post-constitutive equations are well behaved with respect to this perturbation analysis we assert that thermal non-equilibrium should be chosen as an independent force even if this requires the explicit manipulation of the entropy inequality. According to our analysis, alternative choices result in post-constitutive equations; the incompressible limit of which gives rise to questionable predictions.
Electron acceleration in a nonrelativistic shock with very high Alfv\\'en Mach number
Matsumoto, Y; Hoshino, M
2013-01-01
Electron acceleration associated with various plasma kinetic instabilities in a nonrelativistic, very-high-Alfv\\'en Mach-number ($M_A \\sim 45$) shock is revealed by means of a two-dimensional fully kinetic PIC simulation. Electromagnetic (ion Weibel) and electrostatic (ion-acoustic and Buneman) instabilities are strongly activated at the same time in different regions of the two-dimensional shock structure. Relativistic electrons are quickly produced predominantly by the shock surfing mechanism with the Buneman instability at the leading edge of the foot. The energy spectrum has a high-energy tail exceeding the upstream ion kinetic energy accompanying the main thermal population. This gives a favorable condition for the ion acoustic instability at the shock front, which in turn results in additional energization. The large-amplitude ion Weibel instability generates current sheets in the foot, implying another dissipation mechanism via magnetic reconnection in a three-dimensional shock structure in the very-hi...
Capone, F. J.
1972-01-01
An exploratory investigation was conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.20 to 1.30 to determine the induced lift characteristics of a body and swept-wing configuration having a partial-span two-dimensional propulsive nozzle with exhaust exit in the notch of the swept-wing trailing edge. The Reynolds number per meter varied from 4,900,000 to 14,030,000. The effects on wing-body characteristics of deflecting the propulsive jet in the flap mode at nominal exhaust-nozzle deflection angles of 0 deg and 30 deg were studied for two nozzle designs with different geometry and wing spans.
Peng, Naifu; Yang, Yue
2016-11-01
We investigate the evolution of vortex-surface fields (VSFs) in viscous compressible Taylor-Green flows. The VSF is applied to the direct numerical simulation of the Taylor-Green flows at a range of Mach numbers from Ma = 0 . 6 to Ma = 2 . 2 for characterizing the Mach-number effects on evolving vortical structures. We find that the dilatation and baroclinic force strongly influence the geometry of vortex surfaces and the energy dissipation rate in the transitional stage. The vortex tubes in compressible flows are less curved than those in incompressible flows, and the maximum dissipation rate occurs earlier in high-Mach-number flows perhaps owing to the conversion of kinetic energy into heat. Moreover, the relations between the evolutionary geometry of vortical structures and flow statistics are discussed. This work has been supported in part by the National Natural Science Foundation of China (Grant Nos. 11522215 and 11521091), and the Thousand Young Talents Program of China.
Schaeffer, D. B.; Fox, W.; Haberberger, D.; Fiksel, G.; Bhattacharjee, A.; Barnak, D. H.; Hu, S. X.; Germaschewski, K.
2017-07-01
We present the first laboratory generation of high-Mach-number magnetized collisionless shocks created through the interaction of an expanding laser-driven plasma with a magnetized ambient plasma. Time-resolved, two-dimensional imaging of plasma density and magnetic fields shows the formation and evolution of a supercritical shock propagating at magnetosonic Mach number Mms≈12 . Particle-in-cell simulations constrained by experimental data further detail the shock formation and separate dynamics of the multi-ion-species ambient plasma. The results show that the shocks form on time scales as fast as one gyroperiod, aided by the efficient coupling of energy, and the generation of a magnetic barrier between the piston and ambient ions. The development of this experimental platform complements present remote sensing and spacecraft observations, and opens the way for controlled laboratory investigations of high-Mach number collisionless shocks, including the mechanisms and efficiency of particle acceleration.
Artificial viscosity in the transonic stream function formulation
Institute of Scientific and Technical Information of China (English)
徐建中; 杜建一; 沈浩; 刘海涛
1995-01-01
The artificial density method which has been applied widely in the transonic potential calculation and the current transonic stream function calculation is investigated theoretically. The analysis shows that in the stream function formulation the artificial density is not equivalent to the artificial viscosity and cannot be used, and a correct expression of the artificial viscosity in the stream function method is then derived. The principal equation of the stream function, the density equation converted from one of the momentum equations and the present artificial viscosity scheme constitute the complete transonic stream function formulation. The numerical practice demonstrates that the range of Mach number computed by this approach is extended and the shock location is close to the experimental result.
Intermittent Flow Regimes in a Transonic Fan Airfoil Cascade
Lepicovsky, J.; McFarland, E. R.; Chima, R. V.; Capece, V. R.; Hayden, J.
2002-01-01
A study was conducted in the NASA Glenn Research Center linear cascade on the intermittent flow on the suction surface of an airfoil section from the tip region of a modern low aspect ratio fan blade. Experimental results revealed that, at a large incidence angle, a range of transonic inlet Mach numbers exist where the leading-edge shock-wave pattern was unstable. Flush mounted high frequency response pressure transducers indicated large local jumps in the pressure in the leading edge area, which generates large intermittent loading on the blade leading edge. These measurements suggest that for an inlet Mach number between 0.9 and 1.0 the flow is bi-stable, randomly switching between subsonic and supersonic flows. Hence, it appears that the change in overall flow conditions in the transonic region is based on the frequency of switching between two stable flow states rather than on the continuous increase of the flow velocity. To date, this flow behavior has only been observed in a linear transonic cascade. Further research is necessary to confirm this phenomenon occurs in actual transonic fans and is not the byproduct of an endwall restricted linear cascade.
Johnson, F. T.; Samant, S. S.; Bieterman, M. B.; Melvin, R. G.; Young, D. P.; Bussoletti, J. E.; Hilmes, C. L.
1992-01-01
The TranAir computer program calculates transonic flow about arbitrary configurations at subsonic, transonic, and supersonic freestream Mach numbers. TranAir solves the nonlinear full potential equations subject to a variety of boundary conditions modeling wakes, inlets, exhausts, porous walls, and impermeable surfaces. Regions with different total temperature and pressure can be represented. The user's manual describes how to run the TranAir program and its graphical support programs.
Effects of stator bending on pressure field and loss of transonic turbine stage
Institute of Scientific and Technical Information of China (English)
WANG Kai; ZHOU Xun; WANG Zhong-qi
2009-01-01
To study effects of the upstream flow field changing on the downstream flow field of transonic turbine, different three-dimensional bowed blades, which are the stator blades of transonic turbine stage, were designed in this paper. And then numerical calculations were carried out. The effects on downstream flow field were studied and analyzed in detail. Results show that, at the middle of stator blades, although the increasing Mach number causes the increase of shock-wave strength and friction, the middle flow field of downstream rotors is improved obviously. It is an important change in transonic condition. This causes the loss of the rotor's middie part decreased greatly. Correspondingly, efficiency of the whole transonic stage can be increased.
Background-oriented schlieren imaging of flow around a circular cylinder at low Mach numbers
Stadler, Hannes; Bauknecht, André; Siegrist, Silvan; Flesch, Robert; Wolf, C. Christian; van Hinsberg, Nils; Jacobs, Markus
2017-09-01
The background-oriented schlieren (BOS) imaging method has, for the first time, been applied in the investigation of the flow around a circular cylinder at low Mach numbers (Mnumbers of 0.1× 10^6 ≤ Re ≤ 6.0× 10^6. Even at ambient pressure and the lowest Reynolds number investigated, density gradients associated with the flow around the cylinder were recorded. The signal-to-noise ratio of the evaluated gradient field improved with increasing stagnation pressure. The separation point could easily be identified with this non-intrusive measurement technique and corresponds well to simultaneous surface pressure measurements. The resulting displacement field is in principle of qualitative nature as the observation angle was parallel to the cylinder axis only in a single point of the recorded images. However, it has been possible to integrate the density field along the surface of the cylinder by successive imaging at incremental angular positions around the cylinder. This density distribution has been found to agree well with the pressure measurements and with potential theory where appropriate.
Transonic swirling nozzle flow
Keith, Theo G., Jr.; Pawlas, Gary E.
1991-06-01
A numerical model of viscous transonic swirling flow in axisymmetric nozzles is developed. MacCormack's implicit Gauss-Seidel method is applied to the thin-layer Navier-Stokes equations in transformed coordinates. Numerical results are compared with experimental data to validate the method. The effect of swirl and viscosity on nozzle performance are demonstrated by examining wall pressures, Mach contours, and integral parameters.
Aerodynamic Characteristics of a Revised Target Drone Vehicle at Mach Numbers from 1.60 to 2.86
Blair, A. B., Jr.; Babb, C. Donald
1968-01-01
An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a revised target drone vehicle through a Mach number range from 1.60 to 2.86. The vehicle had canard surfaces and a swept clipped-delta wing with twin tip-mounted vertical tails.
Slot, H.J.; Moore, P.; Delfos, R.; Boersma, B.J.
2009-01-01
In this paper we present the experimental results of a detailed investigation of the flow and acoustic properties of a turbulent jet with Mach number 0·75 and Reynolds number 3·5 103. We describe the methods and experimental procedures followed during the measurements, and subsequently present the f
Rescaling of the Roe scheme in low Mach-number flow regions
Boniface, Jean-Christophe
2017-01-01
A rescaled matrix-valued dissipation is reformulated for the Roe scheme in low Mach-number flow regions from a well known family of local low-speed preconditioners popularized by Turkel. The rescaling is obtained explicitly by suppressing the pre-multiplication of the preconditioner with the time derivative and by deriving the full set of eigenspaces of the Roe-Turkel matrix dissipation. This formulation preserves the time consistency and does not require to reformulate the boundary conditions based on the characteristic theory. The dissipation matrix achieves by construction the proper scaling in low-speed flow regions and returns the original Roe scheme at the sonic line. We find that all eigenvalues are nonnegative in the subsonic regime. However, it becomes necessary to formulate a stringent stability condition to the explicit scheme in the low-speed flow regions based on the spectral radius of the rescaled matrix dissipation. With the large disparity of the eigenvalues in the dissipation matrix, this formulation raises a two-timescale problem for the acoustic waves, which is circumvented for a steady-state iterative procedure by the development of a robust implicit characteristic matrix time-stepping scheme. The behaviour of the modified eigenvalues in the incompressible limit and at the sonic line also suggests applying the entropy correction carefully, especially for complex non-linear flows.
Anomalous flow deflection at planetary bow shocks in the low Alfven Mach number regime
Nishino, Masaki N.; Fujimoto, Masaki; Tai, Phan-Duc; Mukai, Toshifumi; Saito, Yoshifumi; Kuznetsova, Masha M.; Rastaetter, Lutz
A planetary magnetosphere is an obstacle to the super-sonic solar wind and the bow shock is formed in the front-side of it. In ordinary hydro-dynamics, the flow decelerated at the shock is diverted around the obstacle symmetrically about the planet-Sun line, which is indeed observed in the magnetosheath most of the time. Here we show a case under a very low density solar wind in which duskward flow was observed in the dawnside magnetosheath of the Earth's magnetosphere. A Rankine-Hugoniot test across the bow shock shows that the magnetic effect is crucial for this "wrong flow" to appear. A full three-dimensional Magneto- Hydro-Dynamics (MHD) simulation of the situation in this previously unexplored parameter regime is also performed. It is illustrated that in addition to the "wrong flow" feature, various peculiar characteristics appear in the global picture of the MHD flow interaction with the obstacle. The magnetic effect at the bow shock should become more conspicuously around the Mercury's magnetosphere, because stronger interplanetary magnetic field and slower solar wind around the Mercury let the Alfven Mach number low. Resultant strong deformation of the magnetosphere induced by the "wrong flow" will cause more complex interaction between the solar wind and the Mercury.
Steady and unsteady transonic flow
Seegmiller, H. L.; Marvin, J. G.; Levy, L. L., Jr.
1978-01-01
An investigation of the transonic flow over a circular arc airfoil was conducted to obtain basic information for turbulence modeling of shock-induced separated flows and to verify numerical computer codes which are being developed to simulate such flows. The investigation included the employment of a laser velocimeter to obtain data concerning the mean velocity, the shear stress, and the turbulent kinetic energy profiles in the flowfield downstream of the airfoil midchord where the flow features are more complex. Depending on the free-stream Mach number, the flowfield developed was either steady with shock-wave-induced separation extending from the foot of the shock wave to beyond the trailing edge or unsteady with a periodic motion also undergoing shock-induced separation. The experimental data were compared with the results of numerical simulations in which a computer code was employed that solved the time-dependent Reynolds' averaged compressible Navier-Stokes equations.
Wallskog, Harvey A.
1954-01-01
A 1/5-scale, rocket-propelled model of the Convair F-102 configuration was tested in free flight to determine zero-lift drag at Mach numbers up to 1.34 and at Reynolds numbers comparable to those of the full-scale airplane. This large-scale model corresponded to the prototype airplane and had air flow through the duct. Additional zero-lift drag tests involved a series of small equivalent bodies of revolution which were launched by means of a helium gun. The several small-scale models tested corresponded to: the basic configuration, the 1/5-scale rocket-propelled model configuration, a 2-foot (full-scale) fuselage-extension configuration, and a 7-foot (full-scale) fuselage-extension configuration. Models designed to correspond to the area distribution at a Mach number of 1.0 were flown for each of these 'shapes and, in addition, models designed to correspond to the area distribution at a Mach number of 1.2 were flown for the 1/5-scale rocket-propelled model and the 7-foot-fuselage-extension configuration. The value of external pressure drag coefficient (including base drag) obtained from the large-scale rocket model was 0.0190 at a Mach number of 1..05 and the corresponding values from the equivalent-body tests varied from 0.0183 for the rocket-propelled model shape to 0.0137 for the 7-foot-fuselage-extension configuration. From the results of tests of equivalent bodies designed to correspond to the area distribution at a Mach number of 1.0, it is evident that the small changes in shape incorporated in the basic and 2-foot-fuselage-extension configurations from that of the rocket-propelled model configuration will provide no significant change in pressure drag. On the other hand, the data from the 7-foot-fuselage-extension model indicate a substantial reduction in pressure drag at transonic speeds.
Pedrosa, A. C. F.; Nagamatsu, H. T.; Hinckel, J. A.
1984-01-01
Heat transfer measurements were determined for a flat plate with and without pressure gradient for various free stream temperatures, wall temperature ratios, and Reynolds numbers for an inlet flow Mach number of 0.45, which is a representative inlet Mach number for gas turbine rotor blades. A shock tube generated the high temperature and pressure air flow, and a variable geometry test section was used to produce inlet flow Mach number of 0.45 and accelerate the flow over the plate to sonic velocity. Thin-film platinum heat gages recorded the local heat flux for laminar, transition, and turbulent boundary layers. The free stream temperatures varied from 611 R (339 K) to 3840 R (2133 K) for a T(w)/T(r,g) temperature ratio of 0.87 to 0.14. The Reynolds number over the heat gages varied from 3000 to 690,000. The experimental heat transfer data were correlated with laminar and turbulent boundary layer theories for the range of temperatures and Reynolds numbers and the transition phenomenon was examined.
Seiff, Alvin; Wilkins, Max E.
1961-01-01
The aerodynamic characteristics of a hypersonic glider configuration, consisting of a slender ogive cylinder with three highly swept wings, spaced 120 apart, with the wing chord equal to the body length, were investigated experimentally at a Mach number of 6 and at Reynolds numbers from 6 to 16 million. The objectives were to evaluate the theoretical procedures which had been used to estimate the performance of the glider, and also to evaluate the characteristics of the glider itself. A principal question concerned the viscous drag at full-scale Reynolds number, there being a large difference between the total drags for laminar and turbulent boundary layers. It was found that the procedures which had been applied for estimating minimum drag, drag due to lift, lift curve slope, and center of pressure were generally accurate within 10 percent. An important exception was the non-linear contribution to the lift coefficient which had been represented by a Newtonian term. Experimentally, the lift curve was nearly linear within the angle-of-attack range up to 10 deg. This error affected the estimated lift-drag ratio. The minimum drag measurements indicated that substantial amounts of turbulent boundary layer were present on all models tested, over a range of surface roughness from 5 microinches maximum to 200 microinches maximum. In fact, the minimum drag coefficients were nearly independent of the surface smoothness and fell between the estimated values for turbulent and laminar boundary layers, but closer to the turbulent value. At the highest test Reynolds numbers and at large angles of attack, there was some indication that the skin friction of the rough models was being increased by the surface roughness. At full-scale Reynolds number, the maximum lift-drag ratio with a leading edge of practical diameter (from the standpoint of leading-edge heating) was 4.0. The configuration was statically and dynamically stable in pitch and yaw, and the center of pressure was less
A comparative study of scramjet injection strategies for high Mach numbers flows
Riggins, D. W.; Mcclinton, C. R.; Rogers, R. C.; Bittner, R. D.
1992-01-01
A simple method for predicting the axial distribution of supersonic combustor thrust potential is described. A complementary technique for illustrating the spatial evolution and distribution of thrust potential and loss mechanisms in reacting flows is developed. Wall jet cases and swept ramp injector cases for Mach 17 and Mach 13.5 flight enthalpy inflow conditions are numerically modeled and analyzed using these techniques. The visualization of thrust potential in the combustor for the various cases examined provides a unique tool for increasing understanding of supersonic combustor performance potential.
Schaeffer, Derek; Haberberger, Dan; Fiksel, Gennady; Bhattacharjee, Amitava; Barnak, Daniel; Hu, Suxing; Germaschewski, Kai
2016-01-01
Shocks act to convert incoming supersonic flows to heat, and in collisionless plasmas the shock layer forms on kinetic plasma scales through collective electromagnetic effects. These collisionless shocks have been observed in many space and astrophysical systems [Smith 1975, Smith 1980, Burlaga 2008, Sulaiman 2015], and are believed to accelerate particles, including cosmic rays, to extremely high energies [Kazanas 1986, Loeb 2000, Bamba 2003, Masters 2013, Ackermann 2013]. Of particular importance are the class of high-Mach number, supercritical shocks [Balogh 2013] ($M_A\\gtrsim4$), which must reflect significant numbers of particles back into the upstream to accommodate entropy production, and in doing so seed proposed particle acceleration mechanisms [Blandford 1978, McClements 2001, Caprioli 2014, Matsumoto 2015]. Here we present the first laboratory generation of high-Mach number magnetized collisionless shocks created through the interaction of an expanding laser-driven plasma with a magnetized ambient ...
Transonic unsteady airloads on an energy efficient transport wing with oscillating control surfaces
Sandford, M. C.; Ricketts, R. H.; Cazier, F. W., Jr.; Cunningham, H. J.
1980-01-01
An aspect ratio 10.8 supercritical wing with oscillating control surfaces is described. The wing is instrumental with 252 static orifices and 164 in situ dynamic pressure transducers for studying the effects of control surface deflection on steady and unsteady pressures at transonic speeds. Results from initial wind tunnel tests conducted in the Langley Transonic Dynamics Tunnel are discussed. Unsteady pressure results are presented for two trailing edge control surfaces oscillating separately at the design Mach number of 0.78. Some experimental results are compared with analytical results obtained by using linear lifting surface theory.
Generation of vertical gusts in a transonic wind tunnel
Brion, V.; Lepage, A.; Amosse, Y.; Soulevant, D.; Senecat, P.; Abart, J. C.; Paillart, P.
2015-07-01
This article reports on the qualification of a gust generator device in a transonic wind tunnel. A vanning apparatus has been installed in the contraction of the S3Ch transonic wind tunnel at the ONERA Meudon center in order to generate up and down air movements in the test section. The apparatus has been tested in a range of Strouhal number based on frequency and vane chord up to 0.15 and in a range of Mach number between 0.3 and 0.73. The amplitude of the gusts has been characterized by a fast-response two-hole pressure probe and phase-averaged PIV. The system delivers vertical velocity amplitude of 0.5 % of the freestream velocity at transonic speeds. For a constant vane oscillation angle, the gust strength is found to increase with the Strouhal and the Mach numbers. The gust exhibit a satisfying uniformity and a quasi-sinusoidal waveform. A simple dynamic point vortex model of the oscillating vanes and of the downstream wake has been developed in order to (1) compare the experimental results and (2) enrich the description of the flow induced by the gusts. In particular, the model is used to analyze the detrimental effect of the upper and lower walls. This simple unsteady model gives a valuable prediction of the amplitude of the gust obtained in the tunnel and the workable frequency range permitted by the present apparatus.
Guo, Xinyi; Sironi, Lorenzo; Narayan, Ramesh
2014-10-01
Electron acceleration to non-thermal energies in low Mach number (Ms Diffusive shock acceleration, also known as first-order Fermi acceleration, cannot be directly invoked to explain the acceleration of electrons. Rather, an additional mechanism is required to pre-accelerate the electrons from thermal to supra-thermal energies, so they can then participate in the Fermi process. In this work, we use two- and three-dimensional particle-in-cell plasma simulations to study electron acceleration in low Mach number shocks. We focus on the particle energy spectra and the acceleration mechanism in a reference run with Ms = 3 and a quasi-perpendicular pre-shock magnetic field. We find that about 15% of the electrons can be efficiently accelerated, forming a non-thermal power-law tail in the energy spectrum with a slope of p ~= 2.4. Initially, thermal electrons are energized at the shock front via shock drift acceleration (SDA). The accelerated electrons are then reflected back upstream where their interaction with the incoming flow generates magnetic waves. In turn, the waves scatter the electrons propagating upstream back toward the shock for further energization via SDA. In summary, the self-generated waves allow for repeated cycles of SDA, similarly to a sustained Fermi-like process. This mechanism offers a natural solution to the conflict between the bright radio synchrotron emission observed from the outskirts of galaxy clusters and the low electron acceleration efficiency usually expected in low Mach number shocks.
Meerson, Baruch; Fouxon, Itzhak; Vilenkin, Arkady
2008-02-01
We employ hydrodynamic equations to investigate nonstationary channel flows of freely cooling dilute gases of hard and smooth spheres with nearly elastic particle collisions. This work focuses on the regime where the sound travel time through the channel is much shorter than the characteristic cooling time of the gas. As a result, the gas pressure rapidly becomes almost homogeneous, while the typical Mach number of the flow drops well below unity. Eliminating the acoustic modes and employing Lagrangian coordinates, we reduce the hydrodynamic equations to a single nonlinear and nonlocal equation of a reaction-diffusion type. This equation describes a broad class of channel flows and, in particular, can follow the development of the clustering instability from a weakly perturbed homogeneous cooling state to strongly nonlinear states. If the heat diffusion is neglected, the reduced equation becomes exactly soluble, and the solution develops a finite-time density blowup. The blowup has the same local features at singularity as those exhibited by the recently found family of exact solutions of the full set of ideal hydrodynamic equations [I. Fouxon, Phys. Rev. E 75, 050301(R) (2007); I. Fouxon,Phys. Fluids 19, 093303 (2007)]. The heat diffusion, however, always becomes important near the attempted singularity. It arrests the density blowup and brings about previously unknown inhomogeneous cooling states (ICSs) of the gas, where the pressure continues to decay with time, while the density profile becomes time-independent. The ICSs represent exact solutions of the full set of granular hydrodynamic equations. Both the density profile of an ICS and the characteristic relaxation time toward it are determined by a single dimensionless parameter L that describes the relative role of the inelastic energy loss and heat diffusion. At L>1 the intermediate cooling dynamics proceeds as a competition between "holes": low-density regions of the gas. This competition resembles Ostwald
Measurement and Analysis of the Noise Radiated by Low Mach Number Centrifugal Blowers.
Yeager, David Marvin
An investigation was performed of the broad band, aerodynamically generated noise in low tip-speed Mach number, centrifugal air moving devices. An interdisciplinary experimental approach was taken which involved investigation of the aerodynamic and acoustic fields, and their mutual relationship. The noise generation process was studied using two experimental vehicles: (1) a scale model of a homologous family of centrifugal blowers typical of those used to cool computer and business equipment, and (2) a single blade from a centrifugal blower impeller placed in a known, controllable flow field. The radiation characteristics of the model blower were investigated by measuring the acoustic intensity distribution near the blower inlet and comparing it with the intensity near the inlet to an axial flow fan. Results showed that the centrifugal blower is a distributed, random noise source, unlike an axial fan which exhibited the effects of a coherent, interacting source distribution. Aerodynamic studies of the flow field in the inlet and at the discharge to the rotating impeller were used to assess the mean flow distribution through the impeller blade channels and to identify regions of excessive turbulence near the rotating blade row. Both circumferential and spanwise mean flow nonuniformities were identified along with a region of increased turbulence just downstream of the scroll cutoff. The fluid incidence angle, normally taken as an indicator of blower performance, was estimated from mean flow data as deviating considerably from an ideal impeller design. An investigation of the noise radiated from the single, isolated airfoil was performed using modern correlation and spectral analysis techniques. Radiation from the single blade in flow was characterized using newly developed expressions for the correlation area and the dipole source strength per unit area, and from the relationship between the blade surface pressure and the incident turbulent flow field. Results
Montoya, L. C.; Jacobs, P. F.; Flechner, S. G.
1977-01-01
Pressure and spanwise load distributions on a first-generation jet transport semispan model at a Mach number of 0.30 are given for the basic wing and for configurations with an upper winglet only, upper and lower winglets, and a simple wing-tip extension. To simulate second-segment-climb lift conditions, leading- and/or trailing-edge flaps were added to some configurations.
Investigation of Three-Dimensional Flow Structure in a Transonic Diffuser by the LIF Method
小野, 大輔; 半田, 太郎; 青木, 俊之; 益田, 光治
2007-01-01
The three-dimensional flow structure induced by normal shock-wave/boundary-layer interaction in a transonic diffuser is investigated by a laser-induced fluorescence (LIF) method. This diagnostic system uses an argon-ion laser as a light source, and target gas is dry nitrogen with iodine seeded as a fluorescence material. The Mach-number distributions in the diffuser are obtained from the measured fluorescence intensity, and the three-dimensional shape of the boundary layers is obtained immedi...
Geared-elevator flutter study. [transonic flutter characteristics of empennage
Ruhlin, C. L.; Doggett, R. V., Jr.; Gregory, R. A.
1976-01-01
The paper describes an experimental and analytical study of the transonic flutter characteristics of an empennage flutter model having an all-movable horizontal tail with a geared elevator. Two configurations were flutter tested: one with a geared elevator and one with a locked elevator with the model cantilever-mounted on a sting in the wind tunnel. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter boundary was nearly flat at transonic speeds for both configurations. It was found that an analysis which treated the elevator as a discrete surface predicted flutter dynamic pressure levels better than analyses which treated the stabilizer and elevator as a warped surface. Warped-surface methods, however, predicted more closely the experimental flutter frequencies and Mach number trends.
Wornom, Dewey E.
1960-01-01
Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.
Seidel, B. S.; Matwey, M. D.; Adamczyk, J. J.
1980-01-01
In the present paper, a semi-actuator-disk theory is reviewed that was developed previously for the distorted inflow to a single-stage axial-flow compressor. Flow distortion occurs far upstream; it may be a distortion in stagnation temperature, stagnation pressure, or both. Losses, quasi-steady deviation angles, and reference incidence correlations are included in the analysis, and both subsonic and transonic relative Mach numbers are considered. The theory is compared with measurements made in a transonic fan stage, and a parameter study is carried out to determine the influence of solidity on the attenuation of distortions in stagnation pressure and stagnation temperature.
Spreading of Exhaust Jet from 16 Inch Ream Jet at Mach Number 2.0 / Fred Wilcox, Donald Pennington
Wilcox, Fred; Pennington, Donald
1952-01-01
An investigation of the jet-spreading characteristics of a 16 inch ram-jet engine was conducted in the 8 by 6 foot supersonic tunnel at a Mach number of 2.0; both a converging nozzle having a contraction ratio of 0.71 and a cylindrical extension to the combustion chamber were used. The jet boundaries determined by means of pitot pressure surveys were compared with boundaries calculated from one-dimensional continuity and momentum relations. For the cylindrical nozzle, the jet reaches its maximum diameter, 4 percent greater than calculated, about 0.6 nozzle-exit diameter downstream of the nozzle exit. The maximum diameter for the converging nozzle was 7 percent greater than calculated from one dimensional relations and occurred from 1 to 1.5 nozzle-exit diameters downstream of the exit. Non dimensional maximum jet diameters agreed closely with results of an investigation by Rousso and Baughman; these data were obtained with low-temperature jets exhausting into a stream at a Mach number of 1.91 from nozzles having exit diameters of 0.75 inch.
Guo, Xinyi; Narayan, Ramesh
2014-01-01
Electron acceleration to non-thermal energies in low Mach number (M<5) shocks is revealed by radio and X-ray observations of galaxy clusters and solar flares, but the electron acceleration mechanism remains poorly understood. Diffusive shock acceleration, also known as first-order Fermi acceleration, cannot be directly invoked to explain the acceleration of electrons. Rather, an additional mechanism is required to pre-accelerate the electrons from thermal to supra-thermal energies, so they can then participate in the Fermi process. In this work, we use two- and three-dimensional particle-in-cell plasma simulations to study electron acceleration in low Mach number shocks. We focus on the particle energy spectra and the acceleration mechanism in a reference run with M=3. We find that about 15 percent of the electrons can be efficiently accelerated, forming a non-thermal power-law tail in the energy spectrum with a slope of p~2.4. Initially, thermal electrons are energized at the shock front via shock drift a...
Syvertson, Clarence A; Gloria, Hermilo R; Sarabia, Michael F
1958-01-01
A study is made of aerodynamic performance and static stability and control at hypersonic speeds. In a first part of the study, the effect of interference lift is investigated by tests of asymmetric models having conical fuselages and arrow plan-form wings. The fuselage of the asymmetric model is located entirely beneath the wing and has a semicircular cross section. The fuselage of the symmetric model was centrally located and has a circular cross section. Results are obtained for Mach numbers from 3 to 12 in part by application of the hypersonic similarity rule. These results show a maximum effect of interference on lift-drag ratio occurring at Mach number of 5, the Mach number at which the asymmetric model was designed to exploit favorable lift interference. At this Mach number, the asymmetric model is indicated to have a lift-drag ratio 11 percent higher than the symmetric model and 15 percent higher than the asymmetric model when inverted. These differences decrease to a few percent at a Mach number of 12. In the course of this part of the study, the accuracy to the hypersonic similarity rule applied to wing-body combinations is demonstrated with experimental results. These results indicate that the rule may prove useful for determining the aerodynamic characteristics of slender configurations at Mach numbers higher than those for which test equipment is really available. In a second part of the study, the aerodynamic performance and static stability and control characteristics of a hypersonic glider are investigated in somewhat greater detail. Results for Mach numbers from 3 to 18 for performance and 0.6 to 12 for stability and control are obtained by standard text techniques, by application of the hypersonic stability rule, and/or by use of helium as a test medium. Lift-drag ratios of about 5 for Mach numbers up to 18 are shown to be obtainable. The glider studied is shown to have acceptable longitudinal and directional stability characteristics through the
Rikanati, A; Oron, D; Sadot, O; Shvarts, D
2003-02-01
Effects of high-Mach numbers and high initial amplitudes on the evolution of the single-mode Richtmyer-Meshkov shock-wave induced hydrodynamic instability are studied using theoretical models, experiments, and numerical simulations. Two regimes in which there is a significant deviation from the linear dependence of the initial velocity on the initial perturbation amplitude are defined and characterized. In one, the observed reduction of the initial velocity is primarily due to large initial amplitudes. This effect is accurately modeled by a vorticity deposition model, quantifying both the effect of the initial perturbation amplitude and the exact shape of the interface. In the other, the reduction is dominated by the proximity of the shock wave to the interface. This effect is modeled by a modified incompressible model where the shock wave is mimicked by a moving bounding wall. These results are supplemented with high initial amplitude Mach 1.2 shock-tube experiments, enabling separation of the two effects. It is shown that in most of the previous experiments, the observed reduction is predominantly due to the effect of high initial amplitudes.
Glazkov, S. A.; Gorbushin, A. R.; Osipova, S. L.; Semenov, A. V.
2016-10-01
The report describes the results of flow field experimental research in TsAGI T-128 transonic wind tunnel. During the tests Mach number, stagnation pressure, test section wall perforation ratio, angles between the test section panels and mixing chamber flaps varied. Based on the test results one determined corrections to the free-stream Mach number related to the flow speed difference in the model location and in the zone of static pressure measurement on the test section walls, nonuniformity of the longitudinal velocity component in the model location, optimal position of the movable test section elements to provide flow field uniformity in the test section and minimize the test leg drag.
Winters, Andrew R.; Derigs, Dominik; Gassner, Gregor J.; Walch, Stefanie
2017-03-01
We describe a unique averaging procedure to design an entropy stable dissipation operator for the ideal magnetohydrodynamic (MHD) and compressible Euler equations. Often in the derivation of an entropy conservative numerical flux function much care is taken in the design and averaging of the entropy conservative numerical flux. We demonstrate in this work that if the discrete dissipation operator is not carefully chosen as well it can have deleterious effects on the numerical approximation. This is particularly true for very strong shocks or high Mach number flows present, for example, in astrophysical simulations. We present the underlying technique of how to construct a unique averaging technique for the discrete dissipation operator. We also demonstrate numerically the increased robustness of the approximation.
Directory of Open Access Journals (Sweden)
Krishna Pandey
2016-01-01
Full Text Available A numerical analysis of the inlet-combustor interaction and flow structure through a scramjet engine at a flight Mach number M = 6 with parallel injection (Strut with circular inlet is presented in the present research article. Three different angles of attack (α=-4°, α=0°, α=4° have been studied for parallel injection. The scramjet configuration used here is a modified version of DLR scramjet model. Fuel is injected at supersonic speed (M=2 through a parallel strut injector. For parallel injection, the shape of the strut is chosen in a way to produce strong stream wise vorticity and thus to enhance the hydrogen/air mixing inside the combustor. These numerical simulations are aimed to study the flow structure, supersonic mixing, and combustion phenomena for the three different types of geometries along with circular shaped strut configuration.
The influence of the Mach number of shock waves on turbulent mixing growth at an interface of gases
Nevmerzhitsky, N. V.; Sotskov, E. A.; Sen'kovsky, E. D.; Razin, A. N.; Ustinenko, V. A.; Krivonos, O. L.; Tochilina, L. V.
2010-12-01
The results of our experimental investigation of the turbulent mixing occurring at a Richtmayer-Meshkov instability driven by a shock wave (SW) in gases at different Mach numbers (M) ranging from ≈1.4 to ≈9 are presented in this paper. The experiments were performed by using an air shock tube with a channel section of 40×40 mm2. The SW passed from 'light' to 'heavy' gases. Air (helium) was used as a 'light' gas and Xe, CO2 and Ar were used as 'heavy' gases. The gases were initially separated by a thin (≈1 μm) polymer film, which was failed after the passing of the SW. A film of the flow was made using a high-speed camera by the Schlieren method.
Wolf, Stephen W. D.; Laub, James A.; King, Lyndell S.; Reda, Daniel C.
1992-01-01
A unique, low-disturbance supersonic wind tunnel is being developed at NASA-Ames to support supersonic laminar flow control research at cruise Mach numbers of the High Speed Civil Transport (HSCT). The distinctive design features of this new quiet tunnel are a low-disturbance settling chamber, laminar boundary layers along the nozzle/test section walls, and steady supersonic diffuser flow. This paper discusses these important aspects of our quiet tunnel design and the studies necessary to support this design. Experimental results from an 1/8th-scale pilot supersonic wind tunnel are presented and discussed in association with theoretical predictions. Natural laminar flow on the test section walls is demonstrated and both settling chamber and supersonic diffuser performance is examined. The full-scale wind tunnel should be commissioned by the end of 1993.
Rao, Pooja; She, Dan; Lim, Hyunkyung; Glimm, James
2015-11-01
The qualitative and quantitative effect of initial conditions (linear and non-linear) and high Mach number (1.3 and 1.45) is studied on the turbulent mixing induced by the Richtmyer-Meshkov instability in idealized ICF conditions. The Richtmyer-Meshkov instability seeds Rayleigh-taylor instabilities in ICF experiments and is one of the factors that contributes to reduced performance of ICF experiments. Its also found in collapsing cores of stars and supersonic combustion. We use the Stony Brook University code, FronTier, which is verified via a code comparison study against the AMR multiphysics code FLASH, and validated against vertical shock tube experiments done by the LANL Extreme Fluids Team. These simulations are designed as a step towards simulating more realistic ICF conditions and quantifying the detrimental effects of mixing on the yield.
Schneider, T.; Botta, N.; Geratz, K. J.; Klein, R.
1999-11-01
When attempting to compute unsteady, variable density flows at very small or zero Mach number using a standard finite volume compressible flow solver one faces at least the following difficulties: (i) Spatial pressure variations vanish as the Mach number M→0, but they do affect the velocity field at leading order; (ii) the resulting spatial homogeneity of the leading order pressure implies an elliptic divergence constraint for the energy flux; (iii) violations of this constraint crucially affect the transport of mass, preventing a code to properly advect even a constant density distribution. We overcome these difficulties through a new algorithm for constructing numerical fluxes in the context of multi-dimensional finite volume methods in conservation form. The construction of numerical fluxes involves: (1) An explicit upwind step yielding predictions for the nonlinear convective flux components. (2) A first correction step that introduces pressure gradients which guarantee compliance of the convective fluxes with a divergence constraint. This step requires the solution of a first Poisson-type equation. (3) A second projection step which provides the yet unknown (non-convective) pressure contribution to the total flux of momentum. This second projection requires the solution of another Poisson-type equation and yields the cell centered velocity field at the new time. This velocity field exactly satisfies a divergence constraint consistent with the asymptotic limit. Step (1) can be done by any standard finite volume compressible flow solver. The input to steps (2) and (3) involves solely the fluxes from step (1) and is independent of how these were obtained. Thus, our approach allows any such solver to be extended to compute variable density incompressible flows.
Unsteady transonic aerodynamics and aeroelastic calculations at low-supersonic freestreams
Guruswamy, Guru P.; Goorjian, Peter M.
1988-01-01
A computational procedure is presented to simulate transonic unsteady flows and corresponding aeroelasticity of wings at low-supersonic freestreams. The flow is modeled by using the transonic small-perturbation theory. The structural equations of motions are modeled using modal equations of motion directly coupled with aerodynamics. Supersonic freestreams are simulated by properly accounting for the boundary conditions based on pressure waves along the flow characteristics in streamwise planes. The flow equations are solved using the time-accurate, alternating-direction implicit finite-difference scheme. The coupled aeroelastic equations of motion are solved by an integration procedure based on the time-accurate, linear-acceleration method. The flow modeling is verified by comparing calculations with experiments for both steady and unsteady flows at supersonic freestreams. The unsteady computations are made for oscillating wings. Comparisons of computed results with experiments show good agreement. Aeroelastic responses are computed for a rectangular wing at Mach numbers ranging from subtransonic to upper-transonic (supersonic) freestreams. The extension of the transonic dip into the upper transonic regime is illustrated.
Marchionna, N. R.; Diehl, L. A.; Trout, A. M.
1973-01-01
Tests were conducted to determine the effect of inlet air humidity on the formation of oxides of nitrogen (NOx) from a gas turbine combustor. Combustor inlet air temperature ranged from 506 K (450 F) to 838 K (1050 F). The tests were primarily run at a constant pressure of 6 atmospheres and reference Mach number of 0.065. The NOx emission index was found to decrease with increasing inlet air humidity at a constant exponential rate: NOx = NOx0e-19H (where H is the humidity and the subscript 0 denotes the value at zero humidity). the emission index increased exponentially with increasing normalized inlet air temperature to the 1.14 power. Additional tests made to determine the effect of pressure and reference Mach number on NOx showed that the NOx emission index varies directly with pressure to the 0.5 power and inversely with reference Mach number.
Coumar, Sandra; Lago, Viviana
2017-06-01
This paper presents an experimental investigation, carried out at the Icare Laboratory by the FAST team, focusing on plasma flow control in supersonic and rarefied regime. The study analyzes how the Mach number as well as the ambient pressure modify the repercussions of the plasma actuator on the shock wave. It follows previous experiments performed in the MARHy (ex-SR3) wind tunnel with a Mach 2 flow interacting with a sharp flat plate, where modifications induced by a plasma actuator were observed. The flat plate was equipped with a plasma actuator composed of two aluminum electrodes. The upstream one was biased with a negative DC potential and thus, created a glow discharge type plasma. Experimental measurements showed that the boundary layer thickness and the shock wave angle increased when the discharge was ignited. The current work was performed with two nozzles generating Mach 4 flows but at two different static pressures: 8 and 71 Pa. These nozzles were chosen to study independently the impact of the Mach number and the impact of the pressure on the flow behavior. In the range of the discharge current considered in this experimental work, it was observed that the shock wave angle increased with the discharge current of +15% for the Mach 2 flow but the increase rate doubled to +28% for the Mach 4 flow at the same static pressure, showing that the discharge effect is even more significant when boosting the flow speed. When studying the effect of the discharge on the Mach 4 flow at higher static pressure, it was observed that the topology of the plasma changed drastically and the increase in the shock wave angle with the discharge current of +21 %.
Practical computational aeroacoustics for compact surfaces in low mach number flows
DEFF Research Database (Denmark)
Pradera-Mallabiabarrena, Ainara; Keith, Graeme; Jacobsen, Finn
2011-01-01
compared to the wavelength of interest. This makes it possible to focus on the surface source term of the Ffowcs Williams-Hawkings equation. In this paper, in order to illustrate the basic method for storing and utilizing data from the CFD analysis, the flow past a circular cylinder at a Reynolds number...
Chan, David T.; Hooker, John R.; Wick, Andrew; Plumley, Ryan W.; Zeune, Cale H.; Ol, Michael V.; DeMoss, Joshua A.
2017-01-01
A wind tunnel investigation of a 0.04-scale model of the Lockheed Martin Hybrid Wing Body (HWB) with Over Wing Nacelles (OWN) air mobility transport configuration was conducted in the National Transonic Facility at the NASA Langley Research Center under a collaborative partnership between NASA, the Air Force Research Laboratory, and Lockheed Martin Aeronautics Company. The wind tunnel test sought to validate the transonic aerodynamic performance of the HWB and to validate the efficiency benefits of the OWN installation as compared to the traditional under-wing installation. The semispan HWB model was tested in a clean wing configuration and also tested with two different nacelles representative of a modern turbofan engine and a future advanced high bypass ratio engine. The nacelles were installed in three different locations with two over-wing positions and one under-wing position. Five-component force and moment data, surface static pressure data, and aeroelastic deformation data were acquired. For the cruise configuration, the model was tested in an angle-of-attack range between -2 and 10 degrees at free-stream Mach numbers from 0.3 to 0.9 and at unit Reynolds numbers between 8 and 39 million per foot, achieving a maximum of 80% of flight Reynolds numbers across the Mach number range. The test results validated pretest computational fluid dynamic (CFD) simulations of the HWB performance including the OWN benefit and the results also exhibited excellent transonic drag data repeatability to within +/-1 drag count. This paper details the experimental setup and model overview, presents some sample data results, and describes the facility improvements that led to the success of the test.
Computational Analysis of the Transonic Dynamics Tunnel Using FUN3D
Energy Technology Data Exchange (ETDEWEB)
Chwalowski, Pawel; Quon, Eliot; Brynildsen, Scott E.
2016-01-04
This paper presents results from an explanatory two-year effort of applying Computational Fluid Dynamics (CFD) to analyze the empty-tunnel flow in the NASA Langley Research Center Transonic Dynamics Tunnel (TDT). The TDT is a continuous-flow, closed circuit, 16- x 16-foot slotted-test-section wind tunnel, with capabilities to use air or heavy gas as a working fluid. In this study, experimental data acquired in the empty tunnel using the R-134a test medium was used to calibrate the computational data. The experimental calibration data includes wall pressures, boundary-layer profiles, and the tunnel centerline Mach number profiles. Subsonic and supersonic flow regimes were considered, focusing on Mach 0.5, 0.7 and Mach 1.1 in the TDT test section. This study discusses the computational domain, boundary conditions, and initial conditions selected in the resulting steady-state analyses using NASA's FUN3D CFD software.
Mack, R. J.
1974-01-01
Wing models were tested in the high-speed section of the Langley Unitary Plan wind tunnel to study the effects of the leading-edge sweep angle and the design lift coefficient on aerodynamic performance and efficiency. The models had leading-edge sweep angles of 69.44 deg, 72.65 deg, and 75.96 deg which correspond to values of the design Mach-number-sweep-angle parameter (beta cotangent A) sub DES of 0.6, 0.75, and 0.9, respectively. For each sweep angle, camber surfaces having design lift coefficients of 0,0.08, and 0.12 at a design Mach number of 2.6 were generated. The wind-tunnel tests were conducted at Mach numbers of 2.3, 2.6, and 2.96 with a stagnation temperature of 338.7 K (150 F) and a Reynolds number per meter of 9.843 times 10 to the 6th power. The results of the tests showed that only a moderate sweeping of the wing leading edge aft of the Mach line along with a small-to-moderate amount of camber and twist was needed to significantly improve the zero-lift (flat camber surface) wing performance and efficiency.
Mahto, Navin Kumar; Choubey, Gautam; Suneetha, Lakka; Pandey, K. M.
2016-11-01
The two equation standard k-ɛ turbulence model and the two-dimensional compressible Reynolds-Averaged Navier-Stokes (RANS) equations have been used to computationally simulate the double cavity scramjet combustor. Here all the simulations are performed by using ANSYS 14-FLUENT code. At the same time, the validation of the present numerical simulation for double cavity has been performed by comparing its result with the available experimental data which is in accordance with the literature. The results are in good agreement with the schlieren image and the pressure distribution curve obtained experimentally. However, the pressure distribution curve obtained numerically is under-predicted in 5 locations by numerical calculation. Further, investigations on the variations of the effects of the length-to-depth ratio of cavity and Mach number on the combustion characteristics has been carried out. The present results show that there is an optimal length-to-depth ratio for the cavity for which the performance of combustor significantly improves and also efficient combustion takes place within the combustor region. Also, the shifting of the location of incident oblique shock took place in the downstream of the H2 inlet when the Mach number value increases. But after achieving a critical Mach number range of 2-2.5, the further increase in Mach number results in lower combustion efficiency which may deteriorate the performance of combustor.
MacArt, Jonathan F.; Mueller, Michael E.
2016-12-01
Two formally second-order accurate, semi-implicit, iterative methods for the solution of scalar transport-reaction equations are developed for Direct Numerical Simulation (DNS) of low Mach number turbulent reacting flows. The first is a monolithic scheme based on a linearly implicit midpoint method utilizing an approximately factorized exact Jacobian of the transport and reaction operators. The second is an operator splitting scheme based on the Strang splitting approach. The accuracy properties of these schemes, as well as their stability, cost, and the effect of chemical mechanism size on relative performance, are assessed in two one-dimensional test configurations comprising an unsteady premixed flame and an unsteady nonpremixed ignition, which have substantially different Damköhler numbers and relative stiffness of transport to chemistry. All schemes demonstrate their formal order of accuracy in the fully-coupled convergence tests. Compared to a (non-)factorized scheme with a diagonal approximation to the chemical Jacobian, the monolithic, factorized scheme using the exact chemical Jacobian is shown to be both more stable and more economical. This is due to an improved convergence rate of the iterative procedure, and the difference between the two schemes in convergence rate grows as the time step increases. The stability properties of the Strang splitting scheme are demonstrated to outpace those of Lie splitting and monolithic schemes in simulations at high Damköhler number; however, in this regime, the monolithic scheme using the approximately factorized exact Jacobian is found to be the most economical at practical CFL numbers. The performance of the schemes is further evaluated in a simulation of a three-dimensional, spatially evolving, turbulent nonpremixed planar jet flame.
Kubiak, M A; Bzowski, M; Sokol, J M; Fuselier, S A; Galli, A; Heirtzler, D; Kucharek, H; Leonard, T W; Moebius, D J McComas E; Park, J; Schwadron, N A; Wurz, P
2016-01-01
With the velocity vector and temperature of the pristine interstellar neutral (ISN) He recently obtained with high precision from a coordinated analysis summarized by McComas et al.2015b, we analyzed the IBEX observations of neutral He left out from this analysis. These observations were collected during the ISN observation seasons 2010---2014 and cover the region in the Earth's orbit where the Warm Breeze persists. We used the same simulation model and a very similar parameter fitting method to that used for the analysis of ISN He. We approximated the parent population of the Warm Breeze in front of the heliosphere with a homogeneous Maxwell-Boltzmann distribution function and found a temperature of $\\sim 9\\,500$ K, an inflow speed of 11.3 km s$^{-1}$, and an inflow longitude and latitude in the J2000 ecliptic coordinates $251.6^\\circ$, $12.0^\\circ$. The abundance of the Warm Breeze relative to the interstellar neutral He is 5.7\\% and the Mach number is 1.97. The newly found inflow direction of the Warm Bree...
Myllys, M. E.; Kilpua, E.; Lavraud, B.
2015-12-01
We have investigated the effect of key solar wind driving parameters on the solar wind-magnetosphere coupling efficiency and saturation of the cross polar cap potential (CPCP) during sheath and magnetic cloud driven storms. The particular focus of the study was on the coupling efficiency dependence with Alfven Mach number (MA).Since we are studying the instantaneous coupling efficiency instead of the average efficiency over the whole solar wind structure, we needed to take into account the communication time between the solar wind and the magnetosphere. We present the results of the time delay analysis between geomagnetic indices (PCN, AE and SYM-H) and the interplanetary electric field y-component (EY, GSM coordinate system) and Newell and Borovsky functions. The study shows that the MA has a clear effect to the saturation of the PCN index, which can be used as a proxy of the polar cap potential. The higher the MA the higher the limit EY value after which the saturation starts to occur. Thus, the coupling efficiency increases as a function of MA. Also, the AE index saturates during high solar wind driving but the saturation is not MA depended. However, the results also suggest that the MA it is not the primary cause for the PCN saturation.
Sovardi, Carlo; Jaensch, Stefan; Polifke, Wolfgang
2016-09-01
A numerical method to concurrently characterize both aeroacoustic scattering and noise sources at a duct singularity is presented. This approach combines Large Eddy Simulation (LES) with techniques of System Identification (SI): In a first step, a highly resolved LES with external broadband acoustic excitation is carried out. Subsequently, time series data extracted from the LES are post-processed by means of SI to model both acoustic propagation and noise generation. The present work studies the aero-acoustic characteristics of an orifice placed in a duct at low flow Mach numbers with the "LES-SI" method. Parametric SI based on the Box-Jenkins mathematical structure is employed, with a prediction error approach that utilizes correlation analysis of the output residuals to avoid overfitting. Uncertainties of model parameters due to the finite length of times series are quantified in terms of confidence intervals. Numerical results for acoustic scattering matrices and power spectral densities of broad-band noise are validated against experimental measurements over a wide range of frequencies below the cut-off frequency of the duct.
Moes, Timothy R.; Whitmore, Stephen A.; Jordan, Frank L., Jr.
1993-01-01
A nonintrusive airdata-sensing system was calibrated in flight and wind-tunnel experiments to an angle of attack of 70 deg and to angles of sideslip of +/- 15 deg. Flight-calibration data have also been obtained to Mach 1.2. The sensor, known as the flush airdata sensor, was installed on the nosecap of an F-18 aircraft for flight tests and on a full-scale F-18 forebody for wind-tunnel tests. Flight tests occurred at the NASA Dryden Flight Research Facility, Edwards, California, using the F-18 High Alpha Research Vehicle. Wind-tunnel tests were conducted in the 30- by 60-ft wind tunnel at the NASA LaRC, Hampton, Virginia. The sensor consisted of 23 flush-mounted pressure ports arranged in concentric circles and located within 1.75 in. of the tip of the nosecap. An overdetermined mathematical model was used to relate the pressure measurements to the local airdata quantities. The mathematical model was based on potential flow over a sphere and was empirically adjusted based on flight and wind-tunnel data. For quasi-steady maneuvering, the mathematical model worked well throughout the subsonic, transonic, and low supersonic flight regimes. The model also worked well throughout the angles-of-attack and -sideslip regions studied.
Tuzla, K.; Russell, D. A.; Wai, J. C.
1976-01-01
Nonlifting 10% biconvex airfoils are mounted in a 30 x 40 cm Ludwieg-tube-driven transonic test-section and the flow field recorded with a holographic interferometer. Nitrogen, argon, and carbon dioxide are used as the principal test gases. Experiments are conducted with Reynolds number based on chord of (0.5-3.5) x 10 to the 6th with Mach numbers of 0.70, 0.75, and 0.80. Supporting calculations use inviscid transonic small-disturbance and full-potential computer codes coupled with simple integral boundary-layer modeling. Systematic studies show that significant gamma-effects can occur due to shock-induced separation.
Eaves, R. H.; Buchanan, T. D.; Warmbrod, J. D.; Johnson, C. B.
1972-01-01
Heat transfer tests for two delta wing configurations were conducted in the hypervelocity wind tunnel. The 24-inch long models were tested at a Mach number of approximately 10.5 and at angles of attack of 20, 40, and 60 degrees over a length Reynolds number range from 5 million to 23 million on 4 May to 4 June 1971. Heat transfer results were obtained from model surface heat gage measurements and thermographic phosphor paint.
Institute of Scientific and Technical Information of China (English)
金志伟; 杨兴锐; 苏北辰
2016-01-01
It is hard to use precise mechanism to describe system dynamic feature of 2.4 m transonic wind tunnel. Put forwards wind tunnel Mach number predictive control strategy based on neural network. Combine the advanteges of model predictive control and nueral network modeling, it is good at processing control parameter unkown, unlinear system and time varing system. Use dynamic response of nueral network based on radial basis function and nonlinear neural network to capture system dynamic feature, apply nerual nwork model in MPC structure. The simulation results show that the control strtegy has a good control effect and trace performance.%针对2.4 m跨声速风洞很难用精确的机理模型表示系统的动态特性的问题,提出了基于神经网络模型的风洞马赫数预测控制策略.综合了模型预测控制和神经网络建模的优点,对于控制参数未知、非线性和时变系统具有很好的处理效果.利用基于径向基函数的神经网络模型预测系统的动态响应、非线性神经网络模型可以在训练过程中捕获系统的动态特性等措施,实现了将神经网络模型应用到MPC结构中.仿真结果表明,该控制策略具有很好的跟踪性能和控制效果.
Contribution from the Earth's Bow Shock to Region 1 Current under Low Alfvén Mach Numbers
Institute of Scientific and Technical Information of China (English)
PENG Zhong; HU You-Qiu
2009-01-01
@@ Using global MHD simulations of the solar wind-magnetosphere--ionosphere system, we investigate the depen-dence of the contribution from the Earth's bow shock (I1bs) to ionospheric region I field aligned current (FAC) (I1). It is found that I1bs increases with increasing southward interplanetary magnetic field (IMF) strength Bs, if the Alfven Mach number MA of the solar wind exceeds 2, a similar result as obtained by previous authors. However, if MA becomes close to or falls below 2, I1bs will decrease with B8 in both magnitude and percentage (i.e., I1bs/I1) because of the resultant reduction of the bow shock strength. Both the surface current density Jbs at the nose of the bow shock and the total bow shock current lb, share nearly the same relationship with MA, and vary non-monotonically with MA or Bs. The maximum point is found to be located at MA = 2.7. Three conclusions are then made as follows: (1) The surface current density at the nose, which is much easier to be evaluated, may be used to largely describe the behaviour of the bow shock instead of the total bow shock current. (2) The peak of the total bow shock current is reached at about MA = 2.7 when only Bs is adjusted. (3) The non-monotonic variation of the bow shock current with MA causes a similar variation of its contribution to region 1 FAC. The turning point for such contribution is found to be nearly MA= 2. The implication of these conclusions to the saturation of the ionospheric transpolar potential is briefly discussed.
Yu, Rixin; Yu, Jiangfei; Bai, Xue-Song
2012-06-01
We present an improved numerical scheme for numerical simulations of low Mach number turbulent reacting flows with detailed chemistry and transport. The method is based on a semi-implicit operator-splitting scheme with a stiff solver for integration of the chemical kinetic rates, developed by Knio et al. [O.M. Knio, H.N. Najm, P.S. Wyckoff, A semi-implicit numerical scheme for reacting flow II. Stiff, operator-split formulation, Journal of Computational Physics 154 (2) (1999) 428-467]. Using the material derivative form of continuity equation, we enhance the scheme to allow for large density ratio in the flow field. The scheme is developed for direct numerical simulation of turbulent reacting flow by employing high-order discretization for the spatial terms. The accuracy of the scheme in space and time is verified by examining the grid/time-step dependency on one-dimensional benchmark cases: a freely propagating premixed flame in an open environment and in an enclosure related to spark-ignition engines. The scheme is then examined in simulations of a two-dimensional laminar flame/vortex-pair interaction. Furthermore, we apply the scheme to direct numerical simulation of a homogeneous charge compression ignition (HCCI) process in an enclosure studied previously in the literature. Satisfactory agreement is found in terms of the overall ignition behavior, local reaction zone structures and statistical quantities. Finally, the scheme is used to study the development of intrinsic flame instabilities in a lean H2/air premixed flame, where it is shown that the spatial and temporary accuracies of numerical schemes can have great impact on the prediction of the sensitive nonlinear evolution process of flame instability.
Park, Jaehong; Workman, Jared C; Blackman, Eric G
2012-01-01
Low Mach number, high beta fast mode shocks can occur in the magnetic reconnection outflows of solar flares. These shocks, which occur above flare loop tops, may provide the electron energization responsible for some of the observed hard X-rays and contemporaneous radio emission. Here we present new 2D particle-in-cell simulations of low Mach number/high beta quasi-perpendicular shocks. The simulations show that electrons above a certain energy threshold experience shock-drift-acceleration. The transition energy between the thermal and non-thermal spectrum and the spectral index from the simulations are consistent with some of the X-ray spectra from RHESSI in the energy regime, $E\\lesssim 40\\sim 100$ keV. Plasma instabilities associated with the shock structure such as the modified-two-stream and the electron whistler/mirror instabilities are examined and compared with the numerical solutions of the kinetic dispersion relations.
Institute of Scientific and Technical Information of China (English)
无
2007-01-01
On super-sonic or trans-sonic planar cascade wind tunnel of free jet intermittent type, wind blowing experiments were performed on the typical sections of stator and rotor blades in the last stage of ultra-ultra-critical steam turbine with extra-long blade of 1200mm. The influences of attack angle and Mach number on the aerodynamic performances of these sections of the blade profiles were verified, and their operating ranges were also specified.
Keenan, James A.; Kuhlman, John M.
1991-01-01
A computational study was conducted on two wings, of aspect ratios 1.244 and 1.865, each having 65 degree leading edge sweep angles, to determine the effects of nonplanar winglets at supersonic Mach numbers. A Mach number of 1.62 was selected as the design value. The winglets studied were parametrically varied in alignment, length, sweep, camber, thickness, and dihedral angle to determine which geometry had the best predicted performance. For the computational analysis, an available Euler marching technique was used. The results indicated that the possibility existed for wing-winglet geometries to equal the performance of wing-alone bodies in supersonic flows with both bodies having the same semispan. The first wing with winglet used NACA 1402 airfoils for the base wing and was shown to have lift-to-pressure drag ratios within 0.136 percent to 0.360 percent of the NACA 1402 wing-alone. The other base wing was a natural flow wing which was previously designed specifically for a Mach number of 1.62. The results obtained showed that the natural wing-alone had a slightly higher lift-to-pressure drag than the natural wing with winglets.
Energy Technology Data Exchange (ETDEWEB)
Li, Pak Shing; Klein, Richard I. [Astronomy Department, University of California, Berkeley, CA 94720 (United States); Martin, Daniel F. [Lawrence Berkeley National Laboratory, 1 Cyclotron Road, Berkeley, CA 94720 (United States); McKee, Christopher F., E-mail: psli@astron.berkeley.edu, E-mail: klein@astron.berkeley.edu, E-mail: DFMartin@lbl.gov, E-mail: cmckee@astro.berkeley.edu [Physics Department and Astronomy Department, University of California, Berkeley, CA 94720 (United States)
2012-02-01
Performing a stable, long-duration simulation of driven MHD turbulence with a high thermal Mach number and a strong initial magnetic field is a challenge to high-order Godunov ideal MHD schemes because of the difficulty in guaranteeing positivity of the density and pressure. We have implemented a robust combination of reconstruction schemes, Riemann solvers, limiters, and constrained transport electromotive force averaging schemes that can meet this challenge, and using this strategy, we have developed a new adaptive mesh refinement (AMR) MHD module of the ORION2 code. We investigate the effects of AMR on several statistical properties of a turbulent ideal MHD system with a thermal Mach number of 10 and a plasma {beta}{sub 0} of 0.1 as initial conditions; our code is shown to be stable for simulations with higher Mach numbers (M{sub rms}= 17.3) and smaller plasma beta ({beta}{sub 0} = 0.0067) as well. Our results show that the quality of the turbulence simulation is generally related to the volume-averaged refinement. Our AMR simulations show that the turbulent dissipation coefficient for supersonic MHD turbulence is about 0.5, in agreement with unigrid simulations.
Li, Pak Shing; Klein, Richard I; McKee, Christopher F
2011-01-01
Performing a stable, long duration simulation of driven MHD turbulence with a high thermal Mach number and a strong initial magnetic field is a challenge to high-order Godunov ideal MHD schemes because of the difficulty in guaranteeing positivity of the density and pressure. We have implemented a robust combination of reconstruction schemes, Riemann solvers, limiters, and Constrained Transport EMF averaging schemes that can meet this challenge, and using this strategy, we have developed a new Adaptive Mesh Refinement (AMR) MHD module of the ORION2 code. We investigate the effects of AMR on several statistical properties of a turbulent ideal MHD system with a thermal Mach number of 10 and a plasma $\\beta_0$ of 0.1 as initial conditions; our code is shown to be stable for simulations with higher Mach numbers ($M_rms = 17.3$) and smaller plasma beta ($\\beta_0 = 0.0067$) as well. Our results show that the quality of the turbulence simulation is generally related to the volume-averaged refinement. Our AMR simulati...
Energy Technology Data Exchange (ETDEWEB)
Kubiak, Marzena A.; Swaczyna, P.; Bzowski, M.; Sokół, J. M. [Space Research Centre of the Polish Academy of Sciences (CBK PAN), 00-716 Warsaw (Poland); Fuselier, S. A.; McComas, D. J. [Southwest Research Institute, San Antonio, TX (United States); Galli, A.; Wurz, P. [Physikalisches Institut, Universität Bern, Bern (Switzerland); Heirtzler, D.; Kucharek, H.; Leonard, T. W.; Möbius, E.; Park, J.; Schwadron, N. A., E-mail: mkubiak@cbk.waw.pl [Space Science Center and Department of Physics, University of New Hampshire, Durham, NH (United States)
2016-04-15
Following the high-precision determination of the velocity vector and temperature of the pristine interstellar neutral (ISN) He via a coordinated analysis summarized by McComas et al., we analyzed the Interstellar Boundary Explorer (IBEX) observations of neutral He left out from this analysis. These observations were collected during the ISN observation seasons 2010–2014 and cover the region in the Earth's orbit where the Warm Breeze (WB) persists. We used the same simulation model and a parameter fitting method very similar to that used for the analysis of ISN He. We approximated the parent population of the WB in front of the heliosphere with a homogeneous Maxwell–Boltzmann distribution function and found a temperature of ∼9500 K, an inflow speed of 11.3 km s{sup −1}, and an inflow longitude and latitude in the J2000 ecliptic coordinates 251.°6, 12.°0. The abundance of the WB relative to ISN He is 5.7% and the Mach number is 1.97. The newly determined inflow direction of the WB, the inflow directions of ISN H and ISN He, and the direction to the center of the IBEX Ribbon are almost perfectly co-planar, and this plane coincides within relatively narrow statistical uncertainties with the plane fitted only to the inflow directions of ISN He, ISN H, and the WB. This co-planarity lends support to the hypothesis that the WB is the secondary population of ISN He and that the center of the Ribbon coincides with the direction of the local interstellar magnetic field (ISMF). The common plane for the direction of the inflow of ISN gas, ISN H, the WB, and the local ISMF is given by the normal direction: ecliptic longitude 349.°7 ± 0.°6 and latitude 35.°7 ± 0.6 in the J2000 coordinates, with a correlation coefficient of 0.85.
Interpolation-based reduced-order modelling for steady transonic flows via manifold learning
DEFF Research Database (Denmark)
Franz, Thomas; Zimmermann, Ralf; Goertz, Stefan
2014-01-01
This paper presents a parametric reduced-order model (ROM) based on manifold learning (ML) for use in steady transonic aerodynamic applications. The main objective of this work is to derive an efficient ROM that exploits the low-dimensional nonlinear solution manifold to ensure an improved...... that has the ability to predict approximate CFD solutions at untried parameter combinations, Isomap is coupled with an interpolation method to capture the variations in parameters like the angle of attack or the Mach number. Furthermore, an approximate local inverse mapping from the reduced...
Parametric study of single expansion ramp nozzles at subsonic/transonic speeds
Capone, F. J.; Re, R. J.; Bare, E. A.; Maclean, M. K.
1987-01-01
The Langley Research Center has conducted a parametric investigation to determine the aeropropulsive characteristics of single expansion ramp nozzles (SERN). The SERN is a nonaxisymmetric, variable-area, internal/external expansion exhaust nozzle. Internal nozzle parameters that were varied included upper ramp length, ramp chordal angle, lower flap length, flap angle and the axial and vertical locations of nozzle throat. Convergent-divergent and convergent nozzles were included in this investigation which was conducted in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.6 to 1.2 and at nozzle pressure ratios up to 12.0.
Pendergraft, O. C., Jr.; Carson, G. T., Jr.
1984-01-01
Static pressure coefficient distributions on the forebody, afterbody, and nozzles of a 1/12 scale F-15 propulsion model was determined in the 16 foot transonic tunnel for Mach numbers from 0.60 to 1.20, angles of attack from -2 deg to 7 deg and ratio of jet total pressure to free stream static pressure from 1 up to about 7, depending on Mach number. The effects of nozzle geometry and horizontal tail deflection on the pressure distributions were investigated. Boundary layer total pressure profiles were determined at two locations ahead of the nozzles on the top nacelle surface. Reynolds number varied from about 1.0 x 10 to the 7th power per meter, depending on Mach number.
Laser anemometer measurements in a transonic axial-flow fan rotor
Strazisar, Anthony J.; Wood, Jerry R.; Hathaway, Michael D.; Suder, Kenneth L.
1989-11-01
Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes.
Laser anemometer measurements in a transonic axial-flow fan rotor
Strazisar, Anthony J.; Wood, Jerry R.; Hathaway, Michael D.; Suder, Kenneth L.
1989-01-01
Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes.
Ng, W.-F.; Rosson, J. C.
1986-01-01
A newly developed, 3-mm-diam, dual hot-wire aspirating probe was used to measure the time-resolved stagnation temperature and pressure in a transonic cryogenic wind tunnel. The probe consists of two coplanar constant temperature hot wires at different overheat ratios operating in a 1.5-mm-diam channel with a choked exit. Thus, the constant Mach number flow by the wires is influenced only by free-stream stagnation temperature and pressure. Diffusion of the free-stream Mach number to a lower value in the channel reduces the dynamic drag on the hot-wire. Frequency response of the present design is dc to 20 kHz. The probe was used to measure the unsteady wake shed from an oscillating airfoil tested in the 0.3-m Transonic Cryogenic Tunnel at NASA-Langley Research Center. The hot-wire lasted for more than ten hours before breaking, proving the ruggedness of the probe and the usefulness of the technique in a high dynamic pressure, transonic cryogenic wind tunnel. Typical data obtained from the experiment are presented after reduction to stagnation pressure and temperature.
3D CFD modeling of subsonic and transonic flowing-gas DPALs with different pumping geometries
Yacoby, Eyal; Sadot, Oren; Barmashenko, Boris D.; Rosenwaks, Salman
2015-10-01
Three-dimensional computational fluid dynamics (3D CFD) modeling of subsonic (Mach number M ~ 0.2) and transonic (M ~ 0.9) diode pumped alkali lasers (DPALs), taking into account fluid dynamics and kinetic processes in the lasing medium is reported. The performance of these lasers is compared with that of supersonic (M ~ 2.7 for Cs and M ~ 2.4 for K) DPALs. The motivation for this study stems from the fact that subsonic and transonic DPALs require much simpler hardware than supersonic ones where supersonic nozzle, diffuser and high power mechanical pump (due to a drop in the gas total pressure in the nozzle) are required for continuous closed cycle operation. For Cs DPALs with 5 x 5 cm2 flow cross section pumped by large cross section (5 x 2 cm2) beam the maximum achievable power of supersonic devices is higher than that of the transonic and subsonic devices by only ~ 3% and ~ 10%, respectively. Thus in this case the supersonic operation mode has no substantial advantage over the transonic one. The main processes limiting the power of Cs supersonic DPALs are saturation of the D2 transition and large ~ 60% losses of alkali atoms due to ionization, whereas the influence of gas heating is negligible. For K transonic DPALs both the gas heating and ionization effects are shown to be unimportant. The maximum values of the power are higher than those in Cs transonic laser by ~ 11%. The power achieved in the supersonic and transonic K DPAL is higher than for the subsonic version, with the same resonator and K density at the inlet, by ~ 84% and ~ 27%, respectively, showing a considerable advantaged of the supersonic device over the transonic one. For pumping by rectangular beams of the same (5 x 2 cm2) cross section, comparison between end-pumping - where the laser beam and pump beam both propagate at along the same axis, and transverse-pumping - where they propagate perpendicularly to each other, shows that the output power and optical-to-optical efficiency are not
Jacobs, P. F.; Flechner, S. G.; Montoya, L. C.
1977-01-01
The effects of winglets and a simple wing-tip extension on the aerodynamic forces and moments and the flow-field cross flow velocity vectors behind the wing tip of a first generation jet transport wing were investigated in the Langley 8-foot transonic pressure tunnel using a semi-span model. The test was conducted at Mach numbers of 0.30, 0.70, 0.75, 0.78, and 0.80. At a Mach number of 0.30, the configurations were tested with combinations of leading- and trailing-edge flaps.
Shrout, B. L.; Corlett, W. A.; Collins, I. K.
1979-01-01
The tabulated results of surface pressure tests conducted on the wing and fuselage of an airplane model in the Langley Unitary Plan wind tunnel are presented without analysis. The model tested was that of a supersonic-cruise airplane with a highly swept arrow-wing planform, two engine nacelles mounted beneath the wing, and outboard vertical tails. Data were obtained at Mach numbers of 2.30, 2.96, and 3.30 for angles of attack from -4 deg to 12 deg. The Reynolds number for these tests was 6,560,000 per meter.
Holland, Scott D.; Murphy, Kelly J.
1993-01-01
Since mission profiles for airbreathing hypersonic vehicles such as the National Aero-Space Plane include single-stage-to-orbit requirements, real gas effects may become important with respect to engine performance. The effects of the decrease in the ratio of specific heats have been investigated in generic three-dimensional sidewall compression scramjet inlets with leading-edge sweep angles of 30 and 70 degrees. The effects of a decrease in ratio of specific heats were seen by comparing data from two facilities in two test gases: in the Langley Mach 6 CF4 Tunnel in tetrafluoromethane (where gamma=1.22) and in the Langley 15-Inch Mach 6 Air Tunnel in perfect gas air (where gamma=1.4). In addition to the simulated real gas effects, the parametric effects of cowl position, contraction ratio, leading-edge sweep, and Reynolds number were investigated in the 15-Inch Mach 6 Air Tunnel. The models were instrumented with a total of 45 static pressure orifices distributed on the sidewalls and baseplate. Surface streamline patterns were examined via oil flow, and schlieren videos were made of the external flow field. The results of these tests have significant implications to ground based testing of inlets in facilities which do not operate at flight enthalpies.
Investigating the Transonic Flutter Boundary of the Benchmark Supercritical Wing
Heeg, Jennifer; Chwalowski, Pawel
2017-01-01
This paper builds on the computational aeroelastic results published previously and generated in support of the second Aeroelastic Prediction Workshop for the NASA Benchmark Supercritical Wing configuration. The computational results are obtained using FUN3D, an unstructured grid Reynolds-Averaged Navier-Stokes solver developed at the NASA Langley Research Center. The analysis results focus on understanding the dip in the transonic flutter boundary at a single Mach number (0.74), exploring an angle of attack range of ??1 to 8 and dynamic pressures from wind off to beyond flutter onset. The rigid analysis results are examined for insights into the behavior of the aeroelastic system. Both static and dynamic aeroelastic simulation results are also examined.
2015-01-02
operate with pulsed NS, PS or combined voltage up to 200 kV (Figure 3). Figure 4. 3D positioning system and Kollmorgen DDR servomotor below the test...section. DDR motor is connected to the model through the 6-component ATI-IA transducer. Pitching mechanism A new pitching mechanism consists of a...to right: UPS power supply; 24-V power supply; NI cRIO 9068 real-time control module; Kollmorgen DDR servomotor interface; main control computer
Energy Technology Data Exchange (ETDEWEB)
Delhaye, D.; Paniagua, G. [von Karman Institute for Fluid Dynamics, Turbomachinery and Propulsion Department, Rhode-Saint-Genese (Belgium); Fernandez Oro, J.M. [Universidad de Oviedo, Area de Mecanica de Fluidos, Gijon (Spain); Denos, R. [European Commission, Directorate General for Research, Brussels (Belgium)
2011-01-15
The paper presents the development and application of a three-sensor wedge probe to measure unsteady aerodynamics in a transonic turbine. CFD has been used to perform a detailed uncertainty analysis related to probe-induced perturbations, in particular the separation zones appearing on the wedge apex. The effects of the Reynolds and Mach numbers are studied using both experimental data together with CFD simulations. The angular range of the probe and linearity of the calibration maps are enhanced with a novel zonal calibration technique, used for the first time in compressible flows. The data reduction methodology is explained and demonstrated with measurements performed in a single-stage high-pressure turbine mounted in the compression tube facility of the von Karman Institute. The turbine was operated at subsonic and transonic pressure ratios (2.4 and 5.1) for a Reynolds number of 10{sup 6}, representative of modern engine conditions. Complete maps of the unsteady flow angle and rotor outlet Mach number are documented. These data allow the study of secondary flows and rotor trailing edge shocks. (orig.)
Delhaye, D.; Paniagua, G.; Fernández Oro, J. M.; Dénos, R.
2011-01-01
The paper presents the development and application of a three-sensor wedge probe to measure unsteady aerodynamics in a transonic turbine. CFD has been used to perform a detailed uncertainty analysis related to probe-induced perturbations, in particular the separation zones appearing on the wedge apex. The effects of the Reynolds and Mach numbers are studied using both experimental data together with CFD simulations. The angular range of the probe and linearity of the calibration maps are enhanced with a novel zonal calibration technique, used for the first time in compressible flows. The data reduction methodology is explained and demonstrated with measurements performed in a single-stage high-pressure turbine mounted in the compression tube facility of the von Karman Institute. The turbine was operated at subsonic and transonic pressure ratios (2.4 and 5.1) for a Reynolds number of 106, representative of modern engine conditions. Complete maps of the unsteady flow angle and rotor outlet Mach number are documented. These data allow the study of secondary flows and rotor trailing edge shocks.
Calleja, John; Tamagno, Jose
1993-01-01
A series of air calibration tests were performed in GASL's HYPULSE facility in order to more accurately determine test section flow conditions for flows simulating total enthalpies in the Mach 13 to 17 range. Present calibration data supplements previous data and includes direct measurement of test section pitot and static pressure, acceleration tube wall pressure and heat transfer, and primary and secondary incident shock velocities. Useful test core diameters along with the corresponding free-stream conditions and usable testing times were determined. For the M13.5 condition, in-stream static pressure surveys showed the temporal and spacial uniformity of this quantity across the useful test core. In addition, finite fringe interferograms taken of the free-stream flow at the test section did not indicate the presence of any 'strong' wave system for any of the conditions investigated.
de Gasperin, F; van Weeren, R J; Dawson, W A; Golovich, N; Wittman, D; Bonafede, A; Bruggen, M
2015-01-01
Diffuse radio emission in the form of radio halos and relics has been found in a number of merging galaxy clusters. These structures indicate that shock and turbulence associated with the merger accelerate electrons to relativistic energies. We report the discovery of a radio relic + radio halo system in PSZ1 G108.18-11.53 (z=0.335). This cluster hosts the second most powerful double radio relic system ever discovered. We observed PSZ1 G108.18-11.53 with the Giant Meterwave Radio Telescope (GMRT) and the Westerbork Synthesis Radio Telescope (WSRT). We obtained radio maps at 147, 323, 607 and 1380 MHz. We also observed the cluster with the Keck telescope, obtaining the spectroscopic redshift for 42 cluster members. From the injection index we obtained the Mach number of the shocks generating the two radio relics. For the southern shock we found M = 2.33^{+0.19}_{-0.26}, while the northern shock Mach number goes from M = 2.20^{+0.07}_{-0.14} in the north part down to M = 2.00^{+0.03}_{-0.08} in the southern reg...
FLEET Velocimetry Measurements on a Transonic Airfoil
Burns, Ross A.; Danehy, Paul M.
2017-01-01
Femtosecond laser electronic excitation tagging (FLEET) velocimetry was used to study the flowfield around a symmetric, transonic airfoil in the NASA Langley 0.3-m TCT facility. A nominal Mach number of 0.85 was investigated with a total pressure of 125 kPa and total temperature of 280 K. Two-components of velocity were measured along vertical profiles at different locations above, below, and aft of the airfoil at angles of attack of 0 deg, 3.5 deg, and 7deg. Measurements were assessed for their accuracy, precision, dynamic range, spatial resolution, and overall measurement uncertainty in the context of the applied flowfield. Measurement precisions as low as 1 m/s were observed, while overall uncertainties ranged from 4 to 5 percent. Velocity profiles within the wake showed sufficient accuracy, precision, and sensitivity to resolve both the mean and fluctuating velocities and general flow physics such as shear layer growth. Evidence of flow separation is found at high angles of attack.
Directory of Open Access Journals (Sweden)
Moritz Schulze
2016-10-01
Full Text Available The interaction of a plane acoustic wave and a sheared flow is numerically investigated for simple orifice and perforated plate configurations in an isolated, non-resonant environment for Mach numbers up to choked conditions in the holes. Analytical derivations found in the literature are not valid in this regime due to restrictions to low Mach numbers and incompressible conditions. To allow for a systematic and detailed parameter study, a low-cost hybrid Computational Fluid Dynamic/Computational Aeroacoustic (CFD/CAA methodology is used. For the CFD simulations, a standard k–ϵ Reynolds-Averaged Navier–Stokes (RANS model is employed, while the CAA simulations are based on frequency space transformed linearized Euler equations (LEE, which are discretized in a stabilized Finite Element method. Simulation times in the order of seconds per frequency allow for a detailed parameter study. From the application of the Multi Microphone Method together with the two-source location procedure, acoustic scattering matrices are calculated and compared to experimental findings showing very good agreement. The scattering properties are presented in the form of scattering matrices for a frequency range of 500–1500 Hz.
Kainulainen, Jouni
2012-01-01
Measuring the mass distribution of infrared dark clouds (IRDCs) over the wide dynamic range of their column densities is a fundamental obstacle in determining the initial conditions of high-mass star formation and star cluster formation. We present a new technique to derive high-dynamic-range, arcsecond-scale resolution column density data for IRDCs and demonstrate the potential of such data in measuring the density variance - sonic Mach number relation in molecular clouds. We combine near-infrared data from the UKIDSS/Galactic Plane Survey with mid-infrared data from the Spitzer/GLIMPSE survey to derive dust extinction maps for a sample of ten IRDCs. We then examine the linewidths of the IRDCs using 13CO line emission data from the FCRAO/Galactic Ring Survey and derive a column density - sonic Mach number relation for them. For comparison, we also examine the relation in a sample of nearby molecular clouds. The presented column density mapping technique provides a very capable, temperature independent tool f...
Main, A. J.; Day, C. R. B.; Lock, G. D.; Oldfield, M. L. G.
1996-08-01
A four-hole pyramid probe has been calibrated for use in a short-duration transonic turbine cascade tunnel. The probe is used to create area traverse maps of total and static pressure, and pitch and yaw angles of the flow downstream of a transonic annular cascade. This data is unusual in that it was acquired in a short-duration (5 s of run time) annular cascade blowdown tunnel. A four-hole pyramid probe was used which has a 2.5 mm section head, and has the side faces inclined at 60° to the flow to improve transonic performance. The probe was calibrated in an ejector driven, perforated wall transonic tunnel over the Mach number range 0.5 1.2, with pitch angles from -20° to + 20° and yaw angles from-23° to +23°. A computer driven automatic traversing mechanism and data collection system was used to acquire a large probe calibration matrix (˜ 10,000 readings) of non dimensional pitch, yaw, Mach number, and total pressure calibration coefficients. A novel method was used to transform the probe calibration matrix of the raw coefficients into a probe application matrix of the physical flow variables (pitch, yaw, Mach number etc.). The probe application matrix is then used as a fast look-up table to process probe results. With negligible loss of accuracy, this method is faster by two orders of magnitude than the alternative of global interpolation on the raw probe calibration matrix. The blowdown tunnel (mean nozzle guide vane blade ring diameter 1.1 m) creates engine representative Reynolds numbers, transonic Mach numbers and high levels (≈ 13%) of inlet turbulence intensity. Contours of experimental measurements at three different engine relevant conditions and two axial positions have been obtained. An analysis of the data is presented which includes a necessary correction for the finite velocity of the probe. Such a correction is non trivial for the case of fast moving probes in compressible flow.
Energy Technology Data Exchange (ETDEWEB)
Core, X.
2002-02-01
The isobar approximation for the system of the balance equations of mass, momentum, energy and chemical species is a suitable approximation to represent low Mach number reactive flows. In this approximation, which neglects acoustics phenomena, the mixture is hydrodynamically incompressible and the thermodynamic effects lead to an uniform compression of the system. We present a novel numerical scheme for this approximation. An incremental projection method, which uses the original form of mass balance equation, discretizes in time the Navier-Stokes equations. Spatial discretization is achieved through a finite volume approach on MAC-type staggered mesh. A higher order de-centered scheme is used to compute the convective fluxes. We associate to this discretization a local mesh refinement method, based on Flux Interface Correction technique. A first application concerns a forced flow with variable density which mimics a combustion problem. The second application is natural convection with first small temperature variations and then beyond the limit of validity of the Boussinesq approximation. Finally, we treat a third application which is a laminar diffusion flame. For each of these test problems, we demonstrate the robustness of the proposed numerical scheme, notably for the density spatial variations. We analyze the gain in accuracy obtained with the local mesh refinement method. (author)
Sandford, M. C.; Ricketts, R. H.; Watson, J. J.
1981-01-01
A high aspect ratio supercritical wing with oscillating control surfaces is described. The semispan wing model was instrumented with 252 static orifices and 164 in situ dynamic pressure gases for studying the effects of control surface position and sinusoidal motion on steady and unsteady pressures. Data from the present test (this is the second in a series of tests on this model) were obtained in the Langley Transonic Dynamics Tunnel at Mach numbers of 0.60 and 0.78 and are presented in tabular form.
Mizukaki, Toshiharu; Borg, Stephen E.; Danehy, Paul M.; Murman, Scott M.
2014-01-01
This paper presents the results of visualization of separated flow around a generic entry capsule that resembles the Apollo Command Module (CM) and the Orion Multi-Purpose Crew Vehicle (MPCV). The model was tested at flow speeds up to Mach 0.4 at a single angle of attack of 28 degrees. For manned spacecraft using capsule-shaped vehicles, certain flight operations such as emergency abort maneuvers soon after launch and flight just prior to parachute deployment during the final stages of entry, the command module may fly at low Mach number. Under these flow conditions, the separated flow generated from the heat-shield surface on both windward and leeward sides of the capsule dominates the wake flow downstream of the capsule. In this paper, flow visualization of the separated flow was conducted using the background-oriented schlieren (BOS) method, which has the capability of visualizing significantly separated wake flows without the particle seeding required by other techniques. Experimental results herein show that BOS has detection capability of density changes on the order of 10(sup-5).
Investigation of Flow Separation in a Transonic-fan Linear Cascade Using Visualization Methods
Lepicovsky, Jan; Chima, Rodrick V.; Jett, Thomas A.; Bencic, Timothy J.; Weiland, Kenneth E.
2000-01-01
An extensive study into the nature of the separated flows on the suction side of modem transonic fan airfoils at high incidence is described in the paper. Suction surface.flow separation is an important flow characteristic that may significantly contribute to stall flutter in transonic fans. Flutter in axial turbomachines is a highly undesirable and dangerous self-excited mode of blade oscillations that can result in high cycle fatigue blade failure. The study basically focused on two visualization techniques: surface flow visualization using dye oils, and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction side of the blade is separated over a large portion of the blade, and the separated area increases with increasing inlet Mach number. For the supersonic inlet flow condition, the flow is attached from the leading edge up to the point where a bow shock from the upper neighboring blade hits the blade surface. Low cascade solidity, for the subsonic inlet flow, results in an increased area of separated flow. For supersonic flow conditions, a low solidity results in an improvement in flow over the suction surface. Finally, computational results modeling the transonic cascade flowfield illustrate our ability to simulate these flows numerically.
Devade, Kiran D.; Pise, Ashok T.
2017-01-01
Ranque Hilsch vortex tube is a device that can produce cold and hot air streams simultaneously from pressurized air. Performance of vortex tube is influenced by a number of geometrical and operational parameters. In this study parametric analysis of vortex tube is carried out. Air is used as the working fluid and geometrical parameters like length to diameter ratio (15, 16, 17, 18), exit valve angles (30°-90°), orifice diameters (5, 6 and 7 mm), 2 entry nozzles and tube divergence angle 4° is used for experimentation. Operational parameters like pressure (200-600 kPa), cold mass fraction (0-1) is varied and effect of Mach number at the inlet of the tube is investigated. The vortex tube is tested at sub sonic (0 tube is observed for CMF up to 0.5. Experimental correlations are proposed for optimum COP. Parametric correlation is developed for geometrical and operational parameters.
Adjoint Method and Predictive Control for 1-D Flow in NASA Ames 11-Foot Transonic Wind Tunnel
Nguyen, Nhan; Ardema, Mark
2006-01-01
This paper describes a modeling method and a new optimal control approach to investigate a Mach number control problem for the NASA Ames 11-Foot Transonic Wind Tunnel. The flow in the wind tunnel is modeled by the 1-D unsteady Euler equations whose boundary conditions prescribe a controlling action by a compressor. The boundary control inputs to the compressor are in turn controlled by a drive motor system and an inlet guide vane system whose dynamics are modeled by ordinary differential equations. The resulting Euler equations are thus coupled to the ordinary differential equations via the boundary conditions. Optimality conditions are established by an adjoint method and are used to develop a model predictive linear-quadratic optimal control for regulating the Mach number due to a test model disturbance during a continuous pitch
Transonic Experimental Research Facility
Federal Laboratory Consortium — The Transonic Experimental Research Facility evaluates aerodynamics and fluid dynamics of projectiles, smart munitions systems, and sub-munitions dispensing systems;...
Lanfranco, M. J.; Sparks, V. W.; Kavanaugh, A. T.
1973-01-01
An experimental investigation was conducted in a 9- by 7-foot supersonic wind tunnel to determine the effect of plume-induced flow separation and aspiration effects due to operation of both the orbiter and the solid rocket motors on a 0.019-scale model of the launch configuration of the space shuttle vehicle. Longitudinal and lateral-directional stability data were obtained at Mach numbers of 1.6, 2.0, and 2.2 with and without the engines operating. The plumes exiting from the engines were simulated by a cold gas jet supplied by an auxiliary 200 atmosphere air supply system, and by solid body plume simulators. Comparisons of the aerodynamic effects produced by these two simulation procedures are presented. The data indicate that the parameters most significantly affected by the jet plumes are the pitching moment, the elevon control effectiveness, the axial force, and the orbiter wing loads.
Miser, James W; Stewart, Warner L
1957-01-01
A blade design study is presented for a two-stage air-cooled turbine suitable for flight at a Mach number of 2.5 for which velocity diagrams have been previously obtained. The detailed procedure used in the design of the blades is given. In addition, the design blade shapes, surface velocity distributions, inner and outer wall contours, and other design data are presented. Of all the blade rows, the first-stage rotor has the highest solidity, with a value of 2.289 at the mean section. The second-stage stator also had a high mean-section solidity of 1.927, mainly because of its high inlet whirl. The second-stage rotor has the highest value of the suction-surface diffusion parameter, with a value of 0.151. All other blade rows have values for this parameter under 0.100.
Voit, Charles H; Guentert, Donald C; Dugan, James F
1950-01-01
A complete stage of an axial-flow compressor was designed and built to investigate the possibility of obtaining a high pressure ratio with an acceptable efficiency through the use of the optimum combination of high blade loading and high relative inlet Mach number. Over-all stage performance was investigated over a range of flows at equivalent tip speeds of 418 to 836 feet per second. At design speed (836 ft/sec), a peak total-pressure ration of 1.445 was obtained with an adiabatic efficiency of 0.89. For design angle of attack at the mean radius, a total-pressure ratio of 1.392 was obtained.
Ruhlin, C. L.; Doggett, R. V., Jr.; Gregory, R. A.
1976-01-01
An experimental and analytical study was made of the transonic flutter characteristics of a supersonic transport tail assembly model having an all-movable, horizontal tail with a geared elevator. Two model configurations, namely, one with a gear-elevator (2.8 to 1.0 gear ratio) and one with locked-elevator (1.0 to 1.0 gear ratio), were flutter tested in the Langley transonic dynamics tunnel with an empennage cantilever-mounted on a sting. The geared-elevator configuration fluttered experimentally at about 20% higher dynamic pressures than the locked-elevator configuration. The experimental flutter dynamic pressure boundaries for both configurations were nearly flat over a Mach number range from 0.9 to 1.1. Flutter calculations (mathematical models) were made for the geared-elevator configuration using three subsonic lifting-surface methods. In one method, the elevator was treated as a discrete surface, and in the other two methods, the stabilizer and elevator were treated as a single warped-surface with the primary difference between these two methods being in the mathematical implementation used. A comparison of the experimental and analytical results shows that the discrete-elevator method predicted best the experimental flutter dynamic pressure level. However, the single warped-surface methods predicts more closely the experimental flutter frequencies and Mach number trends.
Weatherill, W. H.; Sebastian, J. D.; Ehlers, F. E.
1978-01-01
Separating the velocity potential into steady and unsteady parts and linearizing the resulting unsteady equations for small disturbances was performed. The steady velocity potential was obtained first from the well known nonlinear equation for steady transonic flow. The unsteady velocity potential was then obtained from a linear differential equation in complex form with spatially varying coefficients. Since sinusoidal motion is assumed, the unsteady equation is independent of time. The results of an investigation into the relaxation-solution-instability problem was discussed. Concepts examined include variations in outer boundary conditions, a coordinate transformation so that the boundary condition at infinity may be applied to the outer boundaries of the finite difference region, and overlapping subregions. The general conclusion was that only a full direct solution in which all unknowns are obtained at the same time will avoid the solution instabilities of relaxation. An analysis of the one-dimensional form of the unsteady transonic equation was studied to evaluate errors between exact and finite difference solutions. Pressure distributions were presented for a low-aspect-ratio clipped delta wing at Mach number of 0.9 and for a moderate-aspect-ratio rectangular wing at a Mach number of 0.875.
Ruhlin, C. L.; Rauch, F. J., Jr.; Waters, C.
1982-01-01
The model was a 1/6.5-size, semipan version of a wing proposed for an executive-jet-transport airplane. The model was tested with a normal wingtip, a wingtip with winglet, and a normal wingtip ballasted to simulate the winglet mass properties. Flutter and aerodynamic data were acquired at Mach numbers (M) from 0.6 to 0.95. The measured transonic flutter speed boundary for each wingtip configuration had roughly the same shape with a minimum flutter speed near M=0.82. The winglet addition and wingtip mass ballast decreased the wing flutter speed by about 7 and 5 percent, respectively; thus, the winglet effect on flutter was more a mass effect than an aerodynamic effect.
Aerodynamic Measurements of an Incidence Tolerant Blade in a Transonic Turbine Cascade
McVetta, Ashlie B.; Giel, Paul W.
2012-01-01
An overview of the recent facility modifications to NASA s Transonic Turbine Blade Cascade Facility and aerodynamic measurements on the VSPT incidence-tolerant blade are presented. This work supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50% speed range from takeoff to altitude cruise. This results in 50 or more variations in VSPT blade incidence angles. The Transonic Turbine Blade Cascade Facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Details of the modifications are described. An incidence-tolerant blade was developed under an RTPAS study contract and tested in the cascade to look at the effects of large incidence angle and Reynolds number variations. Recent test results are presented which include midspan exit total pressure and flow angle measurements obtained at three inlet angles representing the cruise, take-off, and maximum incidence flight mission points. For each inlet angle, data were obtained at five flow conditions with exit Reynolds numbers varying from 2.12 106 to 2.12 105 and two isentropic exit Mach numbers of 0.72 and 0.35. Three-dimensional flowfield measurements were also acquired at the cruise and take-off points. The flowfield measurements were acquired using a five-hole and three-hole pneumatic probe located in a survey plane 8.6% axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.
Experimental analysis of the shock dynamics on a transonic laminar airfoil
Brion, V.; Dandois, J.; Abart, J.-C.; Paillart, P.
2017-06-01
This paper describes an experimental analysis of the buffet phenomenon on a two-dimensional (2D), transonic, and laminar airfoil at a Reynolds number around 3 · 106. Investigations are carried out in ONERA's S3Ch transonic wind tunnel. The experimental setup allows to vary the Mach number, the angle of attack, and the state of the boundary layer upstream of the shock which can be turbulent or laminar depending on the presence of artificial tripping. Buffet occurs when either the angle of attack or the Mach number is set above a given threshold, which depends upon the particular airfoil, and, as shown here, on the state of the boundary layer. Above the threshold, the boundary layer / shock interaction destabilizes, causing the oscillation of the entire flow field. In the turbulent case, the shock wave moves back and forth over a significant portion of the chord at a frequency of about 75 Hz corresponding to a chord based on Strouhal number St ≃ 0.07, in agreement with previous researches on this phenomenon. In the laminar case, a similar unsteady situation occurs but at a frequency much higher, about 1130 Hz, which corresponds to a Strouhal number of about St ≃ 1. Flow oscillations are limited to the shock foot, the shock itself moving only lightly. The turbulent and laminar bu¨et thresholds are provided. An attempt to apply the classical feedback loop scenario to explain the unsteadiness of the flow in the laminar case is carried out but shows a deceptive agreement with the experimental data. Two other mechanisms of unsteadiness are additionally explored, one based on vortex shedding behind the airfoil and the other on the possible breathing of the laminar separation bubble, which give valuable insights into the §ow physics.
Preliminary results of flow fluctuation measurements in the cryogenic transonic wind tunnel
Zinovyev, V. N.; Lebiga, V. A.; Pak, A. Yu.; Quest, J.
2012-01-01
The detailed information about flow fluctuations structure inside the test section of Pilot of European Transonic Windtunnel (PETW) obtained by means of hot-wire anemometer and fluctuation diagram (FD) method within broad and narrow frequency band is presented. Fluctuation diagrams were derived from an array of hot wire output data measured at different overheating ratio of the probe (not less than 8) at freestream Mach numbers M = 0.2, 0.4, 0.6, 0.7, and 0.8, total temperature T0 = 118 . . . 294.7 K and unit Reynolds numbers Re1 = (5.54 . . . 108.6) · 106 1/m, respectively. Time series of these output signal data were used to obtain information of statistical and correlation features, mode, and spectral composition of flow fluctuations.
Sandford, M. C.; Ricketts, R. H.
1983-01-01
A high aspect ratio supercritical wing with oscillating control surfaces is described. The semispan wing model was instrumented with 252 static pressure orifices and 164 in situ dynamic pressure gages for studying the effects of control surface position and sinusoidal motion on steady and unsteady pressures. Results from the present test (the third in a series of tests on this model) were obtained in the Langley Transonic Dynamics Tunnel at Mach numbers of 0.60, 0.78, and 0.86 and are presented in tabular form.
Chan, David T.; Milholen, William E., II; Jones, Gregory S.; Goodliff, Scott L.
2014-01-01
A second wind tunnel test of the FAST-MAC circulation control semi-span model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged flap. The model was configured for transonic testing of the cruise configuration with 0deg flap deflection to determine the potential for drag reduction with the circulation control blowing. Encouraging results from analysis of wing surface pressures suggested that the circulation control blowing was effective in reducing the transonic drag on the configuration, however this could not be quantified until the thrust generated by the blowing slot was correctly removed from the force and moment balance data. This paper will present the thrust removal methodology used for the FAST-MAC circulation control model and describe the experimental measurements and techniques used to develop the methodology. A discussion on the impact to the force and moment data as a result of removing the thrust from the blowing slot will also be presented for the cruise configuration, where at some Mach and Reynolds number conditions, the thrust-removed corrected data showed that a drag reduction was realized as a consequence of the blowing.
Johnson, D. A.; Rose, W. C.
1976-01-01
Quantitative measurements of the turbulence fluctuations in velocity and mass flux have been obtained in Mach 0.6 and 0.8 turbulent boundary layer and free-shear layer flows by laser velocimetry and hot-wire anemometry techniques. To evaluate the effects of compressibility, these transonic data are compared to available incompressible and supersonic results. Based on some simplifying assumptions, estimates of the rms density fluctuations are made for which error bounds are given. In addition to these fluctuation data, the compressible mean velocity data obtained with the laser velocimeter are presented and compared to pitot tube results. The investigation was conducted in the Ames 6- by 6-Foot Supersonic Wind Tunnel at free-stream Mach numbers of 0.6 and 0.8 for a unit Reynolds number of about 10,000,000 per meter.
Jackson, Charles M., Jr.; Harris, Roy V., Jr.
1960-01-01
An investigation has been made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.99 to determine the longitudinal stability and control characteristics of a reentry model consisting of a lenticular-shaped body with two fin configurations (horizontal fins with end plates). Effects of deflecting the larger size fins as pitch-control surfaces were also investigated. The results indicate that the body alone was unstable from an angle of attack of 0 deg to about 55 deg where it became stable and remained so to 90 deg. The addition of fins provided positive longitudinal stability throughout the angle-of-attack range and increased the lift-drag ratio of the configuration. Reducing the horizontal-fin area at the inboard trailing edge of the fin had only a small effect on the aerodynamic characteristics of the vehicle for the condition of no fin deflection. Deflecting the fins, appeared to be an effective means of pitch control and had only a small effect on lift-drag ratio.
Sayadi, Taraneh; Hamman, Curtis; Moin, Parviz
2011-11-01
Transition to turbulence via spatially evolving secondary instabilities in compressible, zero-pressure-gradient flat plate boundary layers is numerically simulated for both the Klebanoff K-type and Herbert H-type disturbances. The objective of this work is to evaluate the universality of the breakdown process between different routes through transition in wall-bounded shear flows. Each localized linear disturbance is amplified through weak non-linear instability that grows into lambda-vortices and then hairpin-shaped eddies with harmonic wavelength, which become less organized in the late-transitional regime once a fully populated spanwise turbulent energy spectrum is established. For the H-type transition, the computational domain extends from Rex =105 , where laminar blowing and suction excites the most unstable fundamental and a pair of oblique waves, to fully turbulent stage at Rex = 10 . 6 ×105 . The computational domain for the K-type transition extends to Rex = 9 . 6 ×105 . The computational algorithm employs fourth-order central differences with non-reflective numerical sponges along the external boundaries. For each case, the Mach number is 0.2. Supported by the PSAAP program of DoE, ANL and LLNL.
Directory of Open Access Journals (Sweden)
Matthias Bauer
2016-10-01
Full Text Available This paper discusses wind tunnel test results aimed at advancing active flow control technology to increase the aerodynamic efficiency of an aircraft during take-off. A model of the outer section of a representative civil airliner wing was equipped with two-stage fluidic actuators between the slat edge and wing tip, where mechanical high-lift devices fail to integrate. The experiments were conducted at a nominal take-off Mach number of M = 0.2. At this incidence velocity, separation on the wing section, accompanied by increased drag, is triggered by the strong slat edge vortex at high angles of attack. On the basis of global force measurements and local static pressure data, the effect of pulsed blowing on the complex flow is evaluated, considering various momentum coefficients and spanwise distributions of the actuation effort. It is shown that through local intensification of forcing, a momentum coefficient of less than c μ = 0.6 % suffices to offset the stall by 2.4°, increase the maximum lift by more than 10% and reduce the drag by 37% compared to the uncontrolled flow.
Seo, Jung Hee; Mittal, Rajat
2011-02-20
A new sharp-interface immersed boundary method based approach for the computation of low-Mach number flow-induced sound around complex geometries is described. The underlying approach is based on a hydrodynamic/acoustic splitting technique where the incompressible flow is first computed using a second-order accurate immersed boundary solver. This is followed by the computation of sound using the linearized perturbed compressible equations (LPCE). The primary contribution of the current work is the development of a versatile, high-order accurate immersed boundary method for solving the LPCE in complex domains. This new method applies the boundary condition on the immersed boundary to a high-order by combining the ghost-cell approach with a weighted least-squares error method based on a high-order approximating polynomial. The method is validated for canonical acoustic wave scattering and flow-induced noise problems. Applications of this technique to relatively complex cases of practical interest are also presented.
Bzowski, M; Kubiak, M A; Sokol, J M; Fuselier, S A; Galli, A; Heirtzler, D; Kucharek, H; Leonard, T W; McComas, D J; Moebius, E; Schwadron, N A; Wurz, P
2015-01-01
We analyzed observations of interstellar neutral helium (ISN~He) obtained from the Interstellar Boundary Explorer (IBEX) satellite during its first six years of operation. We used a refined version of the ISN~He simulation model, presented in the companion paper by Sokol_et al. 2015, and a sophisticated data correlation and uncertainty system and parameter fitting method, described in the companion paper by Swaczyna et al 2015. We analyzed the entire data set together and the yearly subsets, and found the temperature and velocity vector of ISN~He in front of the heliosphere. As seen in the previous studies, the allowable parameters are highly correlated and form a four-dimensional tube in the parameter space. The inflow longitudes obtained from the yearly data subsets show a spread of ~6 degree, with the other parameters varying accordingly along the parameter tube, and the minimum chi-square value is larger than expected. We found, however, that the Mach number of the ISN~He flow shows very little scatter an...
Introduction to transonic aerodynamics
Vos, Roelof
2015-01-01
Written to teach students the nature of transonic flow and its mathematical foundation, this book offers a much-needed introduction to transonic aerodynamics. The authors present a quantitative and qualitative assessment of subsonic, supersonic, and transonic flow around bodies in two and three dimensions. The book reviews the governing equations and explores their applications and limitations as employed in modeling and computational fluid dynamics. Some concepts, such as shock and expansion theory, are examined from a numerical perspective. Others, including shock-boundary-layer interaction, are discussed from a qualitative point of view. The book includes 60 examples and more than 200 practice problems. The authors also offer analytical methods such as Method of Characteristics (MOC) that allow readers to practice with the subject matter. The result is a wealth of insight into transonic flow phenomena and their impact on aircraft design, including compressibility effects, shock and expansion waves, sho...
Lewis, B. W.
1961-01-01
A limited investigation of the deterioration characteristics of 22 refractory materials was conducted by exposing them to a stagnation temperature of 3,800 F in a Mach number 2 ceramic-heated jet at the Langley Research Center. The materials tested were six materials whose major constituent was silicon carbide, five cermets whose major constituent was titanium carbide, six materials whose major constituents were metal borides, four cermets containing alumina, and one silicon nitride model. Tests consisted of obtaining weight change and appearance changes for 1/2-inch-diameter hemispherical-nose cylindrical models exposed to the air jet for 30 seconds at a time for a total of four runs or 2 minutes exposure. Curves of weight changes plotted against the number of 30-second tests in the jet were obtained. Estimates of average surface temperature near the stagnation point of the model were obtained by use of a special temperature-measuring camera. The models were examined before and after the completion of the tests for possible changes in microstructure; no significant changes were found. The data obtained were analyzed with the view that the oxidation characteristics of the materials were the main factor in deterioration of the materials under the conditions of the tests. It was concluded that only those materials which changed in weight the least could be recommended for further extensive application-oriented evaluations. The following materials fell in this category: silicon carbide - silicon, chromium - 28-percent alumina cermet, titanium boride - 5-percent boron carbide. The remainder of the materials tested had oxidation characteristics which appeared to be too severely limiting of their general applications to flight vehicles.
Swihert, John M
1958-01-01
A brief investigation of a target-type thrust reverser on a single-engine fighter model has been conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.20 to 1.05.At Mach numbers of 0.80, 0.92, and 1.05, a hydrogen peroxide turbojet-engine simulator was operated with the thrust reverser extended. The angle of attack was varied from 0 degrees to 5 degrees at these Mach numbers. The Reynolds number of the free stream, based on the mean aerodynamic chord, was about 5 x 10(6). It was estimated that reversed jet operations separated the model boundary-layer flow over the upper surface of the horizontal tail and upper part of the afterbody. This resulted in a positive pitch increment due to reversed jet operation. Jet-on operation also tended to stabilize the severe lateral oscillations which occurred with the reverser extended and the jet off. It appeared that these jet-off oscillations were the result of an alternating separation and reattachment of the flow on the rearmost portions of the fuselage afterbody.
Gapcynski, John P; Carlson, Harry W
1955-01-01
The changes in the aerodynamic characteristics of a body of revolution with a fineness ratio of 8 have been determined at Mach numbers of 1.41 and 2.01, a Reynolds number, based on body length, of 4.54 x 10 to the 6th power, and angles of incidence of 0 degrees and plus or minus 3 degrees as the position of the body is varied with respect to a reflection plane. The data are compared with theoretical results.
Petronevich, V. V.
2016-10-01
The paper observes the issues related to the increase of efficiency and information content of experimental research in transonic wind tunnels (WT). In particular, questions of optimizing the WT Data Acquisition and Control Systems (DACS) to provide the continuous mode test method are discussed. The problem of Mach number (M number) stabilization in the test section of the large transonic compressor-type wind tunnels at subsonic flow conditions with continuous change of the aircraft model angle of attack is observed on the example of T-128 wind tunnel. To minimize the signals distortion in T-128 DACS measurement channels the optimal MGCplus filter settings of the data acquisition system used in T-128 wind tunnel to measure loads were experimentally determined. As a result of the tests performed a good agreement of the results of balance measurements for pitch/pause and continuous test modes was obtained. Carrying out balance tests for pitch/pause and continuous test methods was provided by the regular data acquisition and control system of T-128 wind tunnel with unified software package POTOK. The architecture and functional abilities of POTOK software package are observed.
Jones, Gregory; Balakrishna, Sundareswara; DeMoss, Joshua; Goodliff, Scott; Bailey, Matthew
2015-01-01
Pressure fluctuations have been measured over the course of several tests in the National Transonic Facility to study unsteady phenomenon both with and without the influence of a model. Broadband spectral analysis will be used to characterize the length scales of the tunnel. Special attention will be given to the large-scale, low frequency data that influences the Mach number and force and moment variability. This paper will also discuss the significance of the vorticity and sound fields that can be related to the Common Research Model and will also highlight the comparisons to an empty tunnel configuration. The effectiveness of vortex generators placed at the interface of the test section and wind tunnel diffuser showed promise in reducing the empty tunnel unsteadiness, however, the vortex generators were ineffective in the presence of a model.
A study of test section configuration for shock tube testing of transonic airfoils
Cook, W. J.
1978-01-01
Two methods are investigated for alleviating wall interference effects in a shock tube test section intended for testing two-dimensional transonic airfoils. The first method involves contouring the test section walls to match approximate streamlines in the flow. Contours are matched to each airfoil tested to produce results close to those obtained in a conventional wind tunnel. Data from a previous study and the present study for two different airfoils demonstrate that useful results are obtained in a shock tube using a test section with contoured walls. The second method involves use of a fixed-geometry slotted-wall test section to provide automatic flow compensation for various airfoils. The slotted-wall test section developed exhibited the desired performance characteristics in the approximate Mach number range 0.82 to 0.89, as evidenced by good agreement obtained between shock tube and wind tunnel results for several airfoil flows.
A design point correlation for losses due to part-span dampers on transonic rotors
Roberts, W. B.
1978-01-01
The design-point losses caused by part-span dampers were correlated for 21 transonic axial-flow fan rotors that had tip speeds varying from 350 to 488 meters per second and design pressure ratios of 1.5 to 2.0. The additional loss attributable to the damper and the total region along the blade height influenced were correlated with selected aerodynamic and geometric parameters. The maximum damper loss correlated well with the mean inlet Mach number at the damper location, the geometric parameters of leading- and trailing-edge damper radius normalized by mean passage height and damper aerodynamic chord, respectively, and the aerodynamic loading parameter of the blade camber divided by the solidity at the damper location. The region of damper influence extended over a mean passage height of the order of 10 to 15 times the maximum damper thickness.
Controlling Compressor Vane Flow Vectoring Angles at Transonic Speeds
Munson, Matthew; Rempfer, Dietmar; Williams, David; Acharya, Mukund
2003-11-01
The ability to control flow separation angles from compressor inlet guide vanes with a Coanda-type actuator is demonstrated using both wind tunnel experiments and finite element simulations. Vectoring angles up to 40 degrees from the uncontrolled baseline state were measured with helium schlieren visualization at transonic Mach numbers ranging from 0.1 to 0.6, and with airfoil chord Reynolds numbers ranging from 89,000 to 710,000. The magnitude of the vectoring angle is shown to depend upon the geometry of the trailing edge, and actuator slot size, and the momentum flux coefficient. Under certain conditions the blowing has no effect on the vectoring angle indicating that the Coanda effect is not present. DNS simulations with the finite element method investigated the effects of geometry changes and external flow. Continuous control of the vectoring angle is demonstrated, which has important implications for application to rotating machinery. The technique is shown to reduce the stall flow coefficient by 15 percent in an axial flow compressor.
Transonic Dynamics Tunnel Force and Pressure Data Acquired on the HSR Rigid Semispan Model
Schuster, David M.; Rausch, Russ D.
1999-01-01
This report describes the aerodynamic data acquired on the High Speed Research Rigid Semispan Model (HSR-RSM) during NASA Langley Transonic Dynamics Tunnel (TDT) Test 520 conducted from 18 March to 4 April, 1996. The purpose of this test was to assess the aerodynamic character of a rigid high speed civil transport wing. The wing was fitted with a single trailing edge control surface which was both steadily deflected and oscillated during the test to investigate the response of the aerodynamic data to steady and unsteady control motion. Angle-of-attack and control surface deflection polars at subsonic, transonic and low-supersonic Mach numbers were obtained in the tunnel?s heavy gas configuration. Unsteady pressure and steady loads data were acquired on the wing, while steady pressures were measured on the fuselage. These data were reduced using a variety of methods, programs and computer systems. The reduced data was ultimately compiled onto a CD-ROM volume which was distributed to HSR industry team members in July, 1996. This report documents the methods used to acquire and reduce the data, and provides an assessment of the quality, repeatability, and overall character of the aerodynamic data measured during this test.
Pierzga, M. J.; Wood, J. R.
1984-01-01
An experimental investigation of the three dimensional flow field through a low aspect ratio, transonic, axial flow fan rotor has been conducted using an advanced laser anemometer (LA) system. Laser velocimeter measurements of the rotor flow field at the design operating speed and over a range of through flow conditions are compared to analytical solutions. The numerical technique used herein yields the solution to the full, three dimensional, unsteady Euler equations using an explicit time marching, finite volume approach. The numerical analysis, when coupled with a simplified boundary layer calculation, generally yields good agreement with the experimental data. The test rotor has an aspect ratio of 1.56, a design total pressure ratio of 1.629 and a tip relative Mach number of 1.38. The high spatial resolution of the LA data matrix (9 radial by 30 axial by 50 blade to blade) permits details of the transonic flow field such as shock location, turning distribution and blade loading levels to be investigated and compared to analytical results.
Pierzga, M. J.; Wood, J. R.
1984-01-01
An experimental investigation of the three-dimensional flow field through a low aspect ratio, transonic, axial flow fan rotor has been conducted, using an advanced laser anemometer (LA) system. Laser velocimeter measurements of the rotor flow field at the design operating speed and over a range of throughflow conditions are compared to analytical solutions. The numerical technique used herein yields the solution to the full, three-dimensional, unsteady Euler equations using an explicit time-marching, finite volume approach. The numerical analysis, when coupled with a simplified boundary layer calculation, generally yields good agreement with the experimental data. The test rotor has an aspect ratio of 1.56, a design total pressure ratio of 1.629 and a tip relative Mach number of 1.38. The high spatial resolution of the LA data matrix (9 radial x 30 axial x 50 blade-to-blade) permits details of the transonic flow field such as shock location, turning distribution, and blade loading levels to be investigated and compared to analytical results.
Investigation of flow separation in a transonic-fan linear cascade using visualization methods
Energy Technology Data Exchange (ETDEWEB)
Lepicovsky, J. [ASRC Aerospace Corporation, NASA GRC, Cleveland, OH (United States)
2008-06-15
An extensive experimental study into the nature of the separated flows on the blade suction surface of modern transonic fans is described in this paper. The study was a subtask of a larger experimental effort focused on blade flutter excited by flow separation in the blade tip region. The tip sections of airfoils on transonic fan blades are designed for precompression and consequently they differ from sections on the rest of the blade. The blade tip section was modeled by a low aspect ratio blade and therefore most of the blade tested was exposed to the secondary flow effects. The aim of this work was to supply reliable data on flow separation on transonic fan blades for validation of future analytical studies. The experimental study focused on two visualization techniques: surface flow visualization using dye oils and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction surface of the blade was separated over a large portion of the blade, and the separated area increased with increasing inlet Mach number. For the supersonic inlet flow condition, the flow was attached from the leading edge up to the point where a bow shock from the upper neighboring blade imposed on the blade surface. Downstream, there was a separated flow region in which air flowed in the direction opposite the inlet flow. Finally, past the separated flow region, the flow reattached to the blade surface. For subsonic inlet flow, the low cascade solidity resulted in an increased area of separated flow. For supersonic flow conditions, the low solidity resulted in an improvement in flow over the suction surface. (orig.)
Investigation of flow separation in a transonic-fan linear cascade using visualization methods
Lepicovsky, J.
2008-06-01
An extensive experimental study into the nature of the separated flows on the blade suction surface of modern transonic fans is described in this paper. The study was a subtask of a larger experimental effort focused on blade flutter excited by flow separation in the blade tip region. The tip sections of airfoils on transonic fan blades are designed for precompression and consequently they differ from sections on the rest of the blade. The blade tip section was modeled by a low aspect ratio blade and therefore most of the blade tested was exposed to the secondary flow effects. The aim of this work was to supply reliable data on flow separation on transonic fan blades for validation of future analytical studies. The experimental study focused on two visualization techniques: surface flow visualization using dye oils and schlieren (and shadowgraph) flow visualization. The following key observations were made during the study. For subsonic inlet flow, the flow on the suction surface of the blade was separated over a large portion of the blade, and the separated area increased with increasing inlet Mach number. For the supersonic inlet flow condition, the flow was attached from the leading edge up to the point where a bow shock from the upper neighboring blade imposed on the blade surface. Downstream, there was a separated flow region in which air flowed in the direction opposite the inlet flow. Finally, past the separated flow region, the flow reattached to the blade surface. For subsonic inlet flow, the low cascade solidity resulted in an increased area of separated flow. For supersonic flow conditions, the low solidity resulted in an improvement in flow over the suction surface.
An exploratory study of a finite difference method for calculating unsteady transonic potential flow
Bennett, R. M.; Bland, S. R.
1979-01-01
A method for calculating transonic flow over steady and oscillating airfoils was developed by Isogai. The full potential equation is solved with a semi-implicit, time-marching, finite difference technique. Steady flow solutions are obtained from time asymptotic solutions for a steady airfoil. Corresponding oscillatory solutions are obtained by initiating an oscillation and marching in time for several cycles until a converged periodic solution is achieved. The method is described in general terms and results for the case of an airfoil with an oscillating flap are presented for Mach numbers 0.500 and 0.875. Although satisfactory results are obtained for some reduced frequencies, it is found that the numerical technique generates spurious oscillations in the indicial response functions and in the variation of the aerodynamic coefficients with reduced frequency. These oscillations are examined with a dynamic data reduction method to evaluate their effects and trends with reduced frequency and Mach number. Further development of the numerical method is needed to eliminate these oscillations.
High-speed flow visualization in hypersonic, transonic, and shock tube flows
Kleine, H.; Olivier, H.
2017-02-01
High-speed flow visualisation has played an important role in the investigations conducted at the Stoßwellenlabor of the RWTH Aachen University for many decades. In addition to applying the techniques of high-speed imaging, this laboratory has been actively developing new or enhanced visualisation techniques and approaches such as various schlieren methods or time-resolved Mach-Zehnder interferometry. The investigated high-speed flows are inherently highly transient, with flow Mach numbers ranging from about M = 0.7 to M = 8. The availability of modern high-speed cameras has allowed us to expand the investigations into problems where reduced reproducibility had so far limited the amount of information that could be extracted from a limited number of flow visualisation records. Following a brief historical overview, some examples of recent studies are given, which represent the breadth of applications in which high-speed imaging has been an essential diagnostic tool to uncover the physics of high-speed flows. Applications include the stability of hypersonic corner flows, the establishment of shock wave systems in transonic airfoil flow, and the complexities of the interactions of shock waves with obstacles of various shapes.
Test facility and the available inlet flow condition to transonic radial diffusers
Energy Technology Data Exchange (ETDEWEB)
Hayami, Hiroshi; Senoo, Yasutoshi; Nakashima, Koji; Kawaguchi, Nobumasa
1987-12-28
In order to stabilize the characteristics combined with diffuser(DF) and expand the flow rate range, a transonic centrifugal impeller(B rotor) with backward blade angle of 40deg was made trially. Single characteristic was tested for five rotating speed equal to or less than 19,000 rpm with freon R12. The flow rate range from choke to inducer(ID) stalling for B rotor is 17% of choke flow rate which is comaratively wider than 11% of conventional impeller(R rotor) with forward blade angle. When the flow field of impeller exit, that is, DF inlet is considered, the maximum value of inlet flow angle in a constant rotating speed corresponds to choke condition or maximum flow rate and minimum value corresponds to operating condition of critical flow rate resulting ID spalling speed. Inflow Mach number to DF changes mainly for the impeller rotating speed. It is possible to pursue the experimental work on diffusers in the range of 21deg(max) to 11.6deg(min) inflow angle and up to 1.31 Mach number using the present facility. (9 figs, 4 refs)
Transonic Investigation of Two-Dimensional Nozzles Designed for Supersonic Cruise
Capone, Francis J.; Deere, Karen A.
2015-01-01
An experimental and computational investigation has been conducted to determine the off-design uninstalled drag characteristics of a two-dimensional convergent-divergent nozzle designed for a supersonic cruise civil transport. The overall objectives were to: (1) determine the effects of nozzle external flap curvature and sidewall boattail variations on boattail drag; (2) develop an experimental data base for 2D nozzles with long divergent flaps and small boattail angles and (3) provide data for correlating computational fluid dynamic predictions of nozzle boattail drag. The experimental investigation was conducted in the Langley 16-Foot Transonic Tunnel at Mach numbers from 0.80 to 1.20 at nozzle pressure ratios up to 9. Three-dimensional simulations of nozzle performance were obtained with the computational fluid dynamics code PAB3D using turbulence closure and nonlinear Reynolds stress modeling. The results of this investigation indicate that excellent correlation between experimental and predicted results was obtained for the nozzle with a moderate amount of boattail curvature. The nozzle with an external flap having a sharp shoulder (no curvature) had the lowest nozzle pressure drag. At a Mach number of 1.2, sidewall pressure drag doubled as sidewall boattail angle was increased from 4deg to 8deg. Reducing the height of the sidewall caused large decreases in both the sidewall and flap pressure drags. Summary
Time-dependent measurement of base pressure in a blowdown tunnel with varying unit Reynolds number
Kangovi, S.; Rao, D. M.
1978-01-01
An operational characteristic of blowdown-type of wind tunnels is the drop in the stagnation temperature with time and the accompanying change in the test-section unit Reynolds number at constant stagnation pressure and Mach number. This apparent disadvantage can be turned to advantage in some cases where a Reynolds number scan is desired in order to study the effect of unit Reynolds number variation on a particular viscous flow phenomenon. This note presents such an instance arising from recent investigations on base pressure at transonic speeds conducted in the NAL 1-ft tunnel.
Harris, C. D.; Bartlett, D. W.
1972-01-01
Basic pressure measurements were made on a 0.087-scale model of a supercritical wing research airplane in the Langley 8 foot transonic pressure tunnel at Mach numbers from 0.25 to 1.00 to determine the effects on the local aerodynamic loads over the wing and rear fuselage of area-rule additions to the sides of the fuselage. In addition, pressure measurements over the surface of the area-rule additions themselves were obtained at angles of sideslip of approximately - 5 deg, 0 deg, and 5 deg to aid in the structural design of the additions. Except for representative figures, results are presented in tabular form without analysis.
Aeroelastic Analyses of the SemiSpan SuperSonic Transport (S4T) Wind Tunnel Model at Mach 0.95
Hur, Jiyoung
2014-01-01
Detailed aeroelastic analyses of the SemiSpan SuperSonic Transport (S4T) wind tunnel model at Mach 0.95 with a 1.75deg fixed angle of attack are presented. First, a numerical procedure using the Computational Fluids Laboratory 3-Dimensional (CFL3D) Version 6.4 flow solver is investigated. The mesh update method for structured multi-block grids was successfully applied to the Navier-Stokes simulations. Second, the steady aerodynamic analyses with a rigid structure of the S4T wind tunnel model are reviewed in transonic flow. Third, the static analyses were performed for both the Euler and Navier-Stokes equations. Both the Euler and Navier-Stokes equations predicted a significant increase of lift forces, compared to the results from the rigid structure of the S4T wind-tunnel model, over various dynamic pressures. Finally, dynamic aeroelastic analyses were performed to investigate the flutter condition of the S4T wind tunnel model at the transonic Mach number. The condition of flutter was observed at a dynamic pressure of approximately 75.0-psf for the Navier-Stokes simulations. However, it was observed that the flutter condition occurred a dynamic pressure of approximately 47.27-psf for the Euler simulations. Also, the computational efficiency of the aeroelastic analyses for the S4T wind tunnel model has been assessed.
Bell, James H.
2011-01-01
The luminescence lifetime technique was used to make pressure-sensitive paint (PSP) measurements on a 2.7% Common Research Model in the NASA Ames 11ft Transonic Wind Tunnel. PSP data were obtained on the upper and lower surfaces of the wing and horizontal tail, as well as one side of the fuselage. Data were taken for several model attitudes of interest at Mach numbers between 0.70 and 0.87. Image data were mapped onto a three-dimensional surface grid suitable both for comparison with CFD and for integration of pressures to determine loads. Luminescence lifetime measurements were made using strobed LED (light-emitting diode) lamps to illuminate the PSP and fast-framing interline transfer cameras to acquire the PSP emission.
Singh, J. J.; Marple, C. G.; Davis, W. T.
1982-01-01
Hot wire anemometry was used to identify the nature of particles reportedly observed during free stream velocity measurements in the Langley 0.3-meter transonic cryogenic tunnel using a Laser Doppler Velocimeter. Since the heat-transfer process from the hot wire depends on the thermal conductivity and sticking capability of the particles, it was anticipated that the hot wire anemometer response would be affected differently upon impaction by liquid droplets and solid aerosols in the test gas stream. Based on the measured time response of the hot-wire anemometer in the cryogenic tunnel operated in the 0.3-0.8 Mach number range, it is concluded that the particles impacting the hot wire are liquid in nature rather than solid aerosols. It is further surmised that the liquid aerosols are unevaporated liquid nitrogen droplets used for cooling the tunnel test gas.
Mitchell, Glenn A; Campbell, Robert C
1957-01-01
Provided sufficient throat bleed was employed, maximum pressure recoveries of 0.87 to 0.88 at Mach number 2.0 were obtained for a fuselage-mounted 14 degrees ramp inlet regardless of the amount of fuselage boundary layer ingested. The addition of inlet side fairings yielded further increases in pressure recovery to 0.90 to 0.91, decreased critical drag coefficients, and increased critical mass-flow ratios. With throat bleed, peak pressure recoveries and calculated thrust-minus-drag values were comparable at two axial positions of the scoop and were highest with the greatest amount of fuselage boundary layer ingested.
Capone, F. J.
1981-01-01
An investigation was conducted in the Langley 16 Foot Transonic Tunnel to determine the aeropropulsive characteristics of a single expansion ramp nozzle (SERN) and a two dimensional convergent divergent nozzle (2-D C-D) installed with both an aft swept and a forward swept wing. The SERN was tested in both an upright and an inverted position. The effects of thrust vectoring at nozzle vector angles from -5 deg to 20 deg were studied. This investigation was conducted at Mach numbers from 0.40 to 1.20 and angles of attack from -2.0 deg to 16 deg. Nozzle pressure ratio was varied from 1.0 (jet off) to about 9.0. Reynolds number based on the wing mean geometric chord varied from about 3 million to 4.8 million, depending upon free stream number.
Ernst Mach a deeper look : documents and new perspectives
1992-01-01
Ernst Mach -- A Deeper Look has been written to reveal to English-speaking readers the recent revival of interest in Ernst Mach in Europe and Japan. The book is a storehouse of new information on Mach as a philosopher, historian, scientist and person, containing a number of biographical and philosophical manuscripts publihsed for the first time, along with correspondence and other matters published for the first time in English. The book also provides English translations of Mach's controversies with leading physicists and psychologists, such as Max Planck and Carl Stumpf, and offers basic evidence for resolving Mach's position on atomism and Einstein's theory of relativity. Mach's scientific, philosophical and personal influence in a number of countries -- Austria, Germany, Bohemia and Yugoslavia among them -- has been carefully explored and many aspects detailed for the first time. All of the articles are eminently readable, especially those written by Mach's sister. They are deeply researched, new interpre...
Institute of Scientific and Technical Information of China (English)
翟永玺; 张堃元; 王磊; 李永洲; 张林
2014-01-01
A parametric research on the curved compression surface with controllable Mach number distri-bution was commenced to find the effect regularity of design parameters on the performance parameters of curved compression surface. On this basis,a polynomial response surface proxy model was built to make a multi-objec-tive optimization,and a hypersonic curved shock two-dimensional inlet was designed based on the optimization result, the performance was compared with the three-ramp compression inlet which was designed under the same constraints. Results indicate among the design parameters, the initial compress angle θ and the factor C and factor md1 affect most. The flow coefficient of the innovative inlet is up to 0.769 at Mach 4,when Mach num-ber ranges from 4 to 7,the two inlets have equally the same mass capture ratio,while the innovative inlet has high total pressure recovery of throat and outlet section. Compared with the relative three-ramp inlet , the total pressure recovery of throat section of the innovative inlet increased by 6.5%at Mach 4, 8.4%at Mach 6, and 10.7%at Mach 7.%针对一种马赫数分布可控的二元高超弯曲压缩面进行参数化研究，获得其设计参数对压缩面性能的影响规律，在此基础上建立多项式响应面代理模型并进行多目标优化，基于优化结果设计了二元弯曲激波进气道，并与同等约束条件下的三楔进气道进行比较。结果表明：压缩面初始压缩角θ与马赫数梯度函数中的设计参数md1，C对压缩面性能影响最为显著；Ma∞=4.0时弯曲激波进气道流量系数达0.769，与三楔进气道相比，在Ma∞=4～7工作范围内的流量捕获能力相当，但其喉道、出口截面的总压恢复系数均高于三楔进气道，在Ma∞=4，6，7工况下，喉道截面总压恢复分别有6.5%，8.4%和10.7%的提高。
Transonic flow theory of airfoils and wings
Energy Technology Data Exchange (ETDEWEB)
Garabedian, P R
1976-01-01
Supercritical wing technology is expected to have a significant influence on the next generation of commercial aircraft. Computational fluid dynamics is playing a central role in the development of new supercritical wing sections. One of the principal tools is a fast and reliable code that simulates two-dimensional wind tunnel data for transonic flow at high Reynolds numbers. This is used widely by industry to assess drag creep and drag rise. Codes for the design of shockless airfoils by the hodograph method have not been so well received because they usually require a lot of trial and error. However, a more advanced mathematical approach makes it possible to assign the pressure as a function of the arc length and then obtain a shockless airfoil that nearly achieves the given distribution of pressure. This tool should enable engineers to design families of transonic airfoils more easily both for airplane wings and for compressor blades in cascade.
Hawthorne, P. J.
1976-01-01
Data obtained in wind tunnel tests are presented. The objectives of the tests were to: (1) obtain pressure distributions, forces and moments over the vehicle 5 Orbiter in the terminal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes. Testing was conducted over a Mach number range from 0.6 to 1.4 with Reynolds number variations from 4.57 million to 2.74 million per foot. Model angle-of-attack was varied from -4 to 16 degrees and angles of side slip ranged from -8 to 8 degrees.
Hawthorne, P. J.
1976-01-01
Data obtained in a wind tunnel test were examined to: (1) obtain pressure distributions, forces and moments over the vehicle 5 Orbiter in the terminal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes. Testing was conducted over a Mach number range from 0.6 to 1.4 with Reynolds number variations from 7.57 x 1 million to 2.74 x 1 million per foot. Model angle of attack was varied from -4 to 16 degrees and angles of sideslip ranged from -8 to 8 degrees.
Hunt, L. Roane; Notestine, Kristopher K.
1990-06-01
Surface and gap pressures and heating-rate distributions were obtained for simulated Thermal Protection System (TPS) tile arrays on the curved surface test apparatus of the Langley 8-Foot High Temperature Tunnel at Mach 6.6. The results indicated that the chine gap pressures varied inversely with gap width because larger gap widths allowed greater venting from the gap to the lower model side pressures. Lower gap pressures caused greater flow ingress from the surface and increased gap heating. Generally, gap heating was greater in the longitudinal gaps than in the circumferential gaps. Gap heating decreased with increasing gap depth. Circumferential gap heating at the mid-depth was generally less than about 10 percent of the external surface value. Gap heating was most severe at local T-gap junctions and tile-to-tile forward-facing steps that caused the greatest heating from flow impingement. The use of flow stoppers at discrete locations reduced heating from flow impingement. The use of flow stoppers at discrete locations reduced heating in most gaps but increased heating in others. Limited use of flow stoppers or gap filler in longitudinal gaps could reduce gap heating in open circumferential gaps in regions of high surface pressure gradients.
Energy Technology Data Exchange (ETDEWEB)
Ansanay-Alex, G.
2009-06-17
The development of simulation codes aimed at a precise simulation of fires requires a precise approach of flame front phenomena by using very fine grids. The need to take different spatial scale into consideration leads to a local grid refinement and to a discretization with homogeneous grid for computing time and memory purposes. The author reports the approximation of the non-linear convection term, the scalar advection-diffusion in finite volumes, numerical simulations of a flow in a bent tube, of a three-dimensional laminar flame and of a low Mach number an-isotherm flow. Non conformal finite elements are also presented (Rannacher-Turek and Crouzeix-Raviart elements)
Jeništa, J.; Takana, H.; Nishiyama, H.; Bartlová, M.; Aubrecht, V.; Křenek, P.; Hrabovský, M.; Kavka, T.; Sember, V.; Mašláni, A.
2011-11-01
This paper presents a numerical investigation of characteristics and processes in the worldwide unique type of thermal plasma generator with combined stabilization of arc by argon flow and water vortex, the so-called hybrid-stabilized arc. The arc has been used for spraying of ceramic or metallic particles and for pyrolysis of biomass. The net emission coefficients as well as the partial characteristics methods for radiation losses from the argon-water arc are employed. Calculations for 300-600 A with 22.5-40 standard litres per minute (slm) of argon reveal transition from a transonic plasma flow for 400 A to a supersonic one for 600 A with a maximum Mach number of 1.6 near the exit nozzle of the plasma torch. A comparison with available experimental data near the exit nozzle shows very good agreement for the radial temperature profiles. Radial velocity profiles calculated 2 mm downstream of the nozzle exit show good agreement with the profiles determined from the combination of calculation and experiment (the so-called integrated approach). A recent evaluation of the Mach number from the experimental data for 500 and 600 A confirmed the existence of the supersonic flow regime.
McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.
2014-01-01
Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50 percent speed range from takeoff to altitude cruise. This results in 50 deg or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83×10 (exp 5) to 0.85×10(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6 percent axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.
Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade
McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.
2013-01-01
Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50 percent speed range from takeoff to altitude cruise. This results in 50deg or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83×10(exp 5) to 0.85×10(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6 percent axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.
Hawthorne, P. J.
1976-01-01
Data obtained in wind tunnel test OA148 are presented. The objectives of the test series were to: (1) obtain pressure distributions, forces and moments over the vehicle 5 orbiter in the thermal area energy management (TAEM) and approach phases of flight; (2) obtain elevon and rudder hinge moments in the TAEM and approach phases of flight; (3) obtain body flap and elevon loads for verification of loads balancing with integrated pressure distributions; and (4) obtain pressure distributions near the short OMS pods in the high subsonic, transonic and low supersonic Mach number regimes.
VTOL to Transonic Aircraft Project
National Aeronautics and Space Administration — The cyclogyro, an aircraft propulsion concept with the potential for VTOL to the lower bounds of transonic flight, is conceptually simple but structurally and...
Tavelli, Maurizio; Dumbser, Michael
2017-07-01
We propose a new arbitrary high order accurate semi-implicit space-time discontinuous Galerkin (DG) method for the solution of the two and three dimensional compressible Euler and Navier-Stokes equations on staggered unstructured curved meshes. The method is pressure-based and semi-implicit and is able to deal with all Mach number flows. The new DG scheme extends the seminal ideas outlined in [1], where a second order semi-implicit finite volume method for the solution of the compressible Navier-Stokes equations with a general equation of state was introduced on staggered Cartesian grids. Regarding the high order extension we follow [2], where a staggered space-time DG scheme for the incompressible Navier-Stokes equations was presented. In our scheme, the discrete pressure is defined on the primal grid, while the discrete velocity field and the density are defined on a face-based staggered dual grid. Then, the mass conservation equation, as well as the nonlinear convective terms in the momentum equation and the transport of kinetic energy in the energy equation are discretized explicitly, while the pressure terms appearing in the momentum and energy equation are discretized implicitly. Formal substitution of the discrete momentum equation into the total energy conservation equation yields a linear system for only one unknown, namely the scalar pressure. Here the equation of state is assumed linear with respect to the pressure. The enthalpy and the kinetic energy are taken explicitly and are then updated using a simple Picard procedure. Thanks to the use of a staggered grid, the final pressure system is a very sparse block five-point system for three dimensional problems and it is a block four-point system in the two dimensional case. Furthermore, for high order in space and piecewise constant polynomials in time, the system is observed to be symmetric and positive definite. This allows to use fast linear solvers such as the conjugate gradient (CG) method. In
Phillips, W. P.
1984-01-01
Aerodynamic characteristics at M=5.97 for the 140 A/B Space Shuttle Orbiter configuration and for the configuration modified by geometric changes in the wing planform fillet region and the fuselage forebody are presented. The modifications, designed to extend the orbiter's longitudinal trim capability to more forward center of gravity locations, include reshaping the baseline wing fillet, changing the fuselage forebody camber, and adding canards. The Langley 20 inch Mach 6 Tunnel at a Reynolds number of approximately 6 million based on fuselage reference length was used. The angle of attack range of the investigation varied from about 15 deg to 35 deg at 0 deg and -5 deg sideslip angles. Data are obtained with the elevators and body flap deflected at appropriate negative and positive conditions to assess the trim limits.
Energy Technology Data Exchange (ETDEWEB)
Karasawa, T.; Suenaga, H.; Suzuki, M.; Komatsu, Y.; Toda, N. [National Aerospace Laboratory, Tokyo (Japan)
1996-08-01
This paper describes an outline of a replaced auxiliary blower system for NAL 2m{times}2m transonic wind tunnel, and effects of its renewal. Based on the guideline of renewal of facilities, the bleed air flow was controlled by the whole step stator blade control method of main body of the blower, and a blower with a capacity adjustment function was equipped. Downsizing was planned by adopting an accelerator and a drive motor having their rated maximum power of 8,000 kW, which increased the number of revolution to 1.5 times. The protective function for antisurging control of the blower was enhanced by the by-pass control valve, which provided the auxiliary function for the Mach number control. The new valve was downsized to 6/7 in the bore diameter due to the reduction of by-pass air flow compared with the previous valve. Results of the present renewal are as follows. By the function recovery, tests in the Mach number up to 1.4 can be conducted under the total pressure of collecting drum, 120 kPa. Furthermore, energy saving, automatic control of the Mach number, operation and monitoring by the centralized processing, and high efficiency of operation were achieved. 6 refs., 25 figs., 1 tab.
Study on Mach stems induced by interaction of planar shock waves on two intersecting wedges
Institute of Scientific and Technical Information of China (English)
Gaoxiang Xiang; Chun Wang; Honghui Teng; Yang Yang; Zonglin Jiang
2016-01-01
The properties of Mach stems in hypersonic corner flow induced by Mach interaction over 3D inter-secting wedges were studied theoretically and numerically. A new method called “spatial dimension reduction” was used to analyze theoretically the location and Mach num-ber behind Mach stems. By using this approach, the problem of 3D steady shock/shock interaction over 3D intersecting wedges was transformed into a 2D moving one on cross sec-tions, which can be solved by shock-polar theory and shock dynamics theory. The properties of Mach interaction over 3D intersecting wedges can be analyzed with the new method, including pressure, temperature, density in the vicinity of triple points, location, and Mach number behind Mach stems. Theoretical results were compared with numerical results, and good agreement was obtained. Also, the influence of Mach number and wedge angle on the properties of a 3D Mach stem was studied.
Pfyl, Frank A.; Presley, Leroy L.
1961-01-01
The local recovery factor was determined experimentally along the surface of a thin-walled 20 deg included angle cone for Mach numbers near 6.0 at stagnation temperatures between 1200 deg R and 2600 deg R. In addition, a similar cone configuration was tested at Mach numbers near 4.5 at stagnation temperatures of approximately 612 deg R. The local Reynolds number based on flow properties at the edge of the boundary layer ranged between 0.1 x 10(exp 4) and 3.5 x 10(exp 4) for tests at temperatures above 1200 deg R and between 6 x 10(exp 4) and 25 x 10(exp 4) for tests at temperatures near 612 deg R. The results indicated, generally, that the recovery factor can be predicted satisfactorily using the square root of the Prandtl number. No conclusion could be made as to the necessity of evaluating the Prandtl number at a reference temperature given by an empirical equation, as opposed to evaluating the Prandtl number at the wall temperature or static temperature of the gas at the cone surface. For the tests at temperatures above 1200 deg R (indicated herein as the tests conducted in the slip-flow region), two definite trends in the recovery data were observed - one of increasing recovery factor with decreasing stagnation pressure, which was associated with slip-flow effects and one of decreasing recovery factor with increasing temperature. The true cause of the latter trend could not be ascertained, but it was shown that this trend was not appreciably altered by the sources of error of the magnitude considered herein. The real-gas equations of state were used to determine accurately the local stream properties at the outer edge of the boundary layer of the cone. Included in the report, therefore, is a general solution for the conical flow of a real gas using the Beattie-Bridgeman equation of state. The largest effect of temperature was seen to be in the terms which were dependent upon the internal energy of the gas. The pressure and hence the pressure drag terms were
Bhatia, K. G.; Nagaraja, K. S.
1984-01-01
Flutter characteristics of a cantilevered high aspect ratio wing with winglet were investigated. The configuration represented a current technology, twin engine airplane. Compressibility effects through transonic Mach numbers and a wide range of mass-density ratios were evaluated on a low speed and high speed model. Four flutter mechanisms were obtained from test, and analysis from various combinations of configuration parameters. It is shown that the coupling between wing tip vertical and chordwise motions have significant effect under some conditions. It is concluded that for the flutter model configurations studied, the winglet related flutter is amenable to the conventional flutter analysis techniques. The low speed model flutter and the high-speed model flutter results are described.
McMIllin, S. Naomi; Byrd, James E.; Parmar, Devendra S.; Bezos-O'Connor, Gaudy M.; Forrest, Dana K.; Bowen, Susan
1996-01-01
An experimental investigation of the effect of leading-edge radius, camber, Reynolds number, and boundary-layer state on the incipient separation of a delta wing at supersonic speeds was conducted at the Langley Unitary Plan Wind Tunnel at Mach number of 1.60 over a free-stream Reynolds number range of 1 x 106 to 5 x 106 ft-1. The three delta wing models examined had a 65 deg swept leading edge and varied in cross-sectional shape: a sharp wedge, a 20:1 ellipse, and a 20:1 ellipse with a -9.750 circular camber imposed across the span. The wings were tested with and without transition grit applied. Surface-pressure coefficient data and flow-visualization data are electronically stored on the CD-ROM. The data indicated that by rounding the wing leading edge or cambering the wing in the spanwise direction, the onset of leading-edge separation on a delta wing can be raised to a higher angle of attack than that observed on a sharp-edged delta wing. The data also showed that the onset of leading-edge separation can be raised to a higher angle of attack by forcing boundary-layer transition to occur closer to the wing leading edge by the application of grit or the increase in free-stream Reynolds number.
Evaluation of turbulence models in the PARC code for transonic diffuser flows
Georgiadis, N. J.; Drummond, J. E.; Leonard, B. P.
1994-01-01
Flows through a transonic diffuser were investigated with the PARC code using five turbulence models to determine the effects of turbulence model selection on flow prediction. Three of the turbulence models were algebraic models: Thomas (the standard algebraic turbulence model in PARC), Baldwin-Lomax, and Modified Mixing Length-Thomas (MMLT). The other two models were the low Reynolds number k-epsilon models of Chien and Speziale. Three diffuser flows, referred to as the no-shock, weak-shock, and strong-shock cases, were calculated with each model to conduct the evaluation. Pressure distributions, velocity profiles, locations of shocks, and maximum Mach numbers in the duct were the flow quantities compared. Overall, the Chien k-epsilon model was the most accurate of the five models when considering results obtained for all three cases. However, the MMLT model provided solutions as accurate as the Chien model for the no-shock and the weak-shock cases, at a substantially lower computational cost (measured in CPU time required to obtain converged solutions). The strong shock flow, which included a region of shock-induced flow separation, was only predicted well by the two k-epsilon models.
Brown, C. A., Jr.; Campbell, J. F.; Tudor, D. H.
1971-01-01
An investigation was conducted to obtain flow properties in the wake of the Viking '75 entry vehicle at Mach numbers from 1.60 to 3.95 and at angles of attack of 0 deg and 5 deg. The wake flow properties were calculated from total and static pressures measured with a pressure rake at longitudinal stations varying from 1.0 to 8.39 body diameters and lateral stations varying from -0.42 to 3.0 body diameters. These measurements showed a a consistent trend throughout the range of Mach numbers and longitudinal distances and an increase in dynamic pressure with increasing downstream position.
Flow Control in a Transonic Diffuser
Gartner, Jeremy; Amitay, Michael
2014-11-01
In some airplanes such as fighter jets and UAV, short inlet ducts replace the more conventional ducts due to their shorter length. However, these ducts are associated with low length-to-diameter ratio and low aspect ratio and, thus, experience massive separation and the presence of secondary flow structures. These flow phenomena are undesirable as they lead to pressure losses and distortion at the Aerodynamic Interface Plane (AIP), where the engine face is located. It causes the engine to perform with a lower efficiency as it would with a straight duct diffuser. Different flow control techniques were studied on the short inlet duct, with the goal to reattach the flow and minimize the distortions at the AIP. Due to the complex interaction between the separation and the secondary flow structures, the necessity to understand the flow mechanisms, and how to control them at a more fundamental level, a new transonic diffuser with an upper ramp and a straight floor was designed and built. The objective of this project is to explore the effectiveness of different flow control techniques in a high subsonic (up to Mach 0.8) diffuser, so that the quasi two-dimensional separation and the formation of secondary flow structure can be isolated using a canonical flow field. Supported by Northrop Grumman.
Jump conditions in transonic equilibria
Energy Technology Data Exchange (ETDEWEB)
Guazzotto, L.; Betti, R. [Department of Mechanical Engineering, University of Rochester, Rochester, New York 14627 (United States); Jardin, S. C. [Princeton Plasma Physics Laboratory, Princeton, New Jersey 08540 (United States)
2013-04-15
In the present paper, the numerical calculation of transonic equilibria, first introduced with the FLOW code in Guazzotto et al.[Phys. Plasmas 11, 604 (2004)], is critically reviewed. In particular, the necessity and effect of imposing explicit jump conditions at the transonic discontinuity are investigated. It is found that 'standard' (low-{beta}, large aspect ratio) transonic equilibria satisfy the correct jump condition with very good approximation even if the jump condition is not explicitly imposed. On the other hand, it is also found that high-{beta}, low aspect ratio equilibria require the correct jump condition to be explicitly imposed. Various numerical approaches are described to modify FLOW to include the jump condition. It is proved that the new methods converge to the correct solution even in extreme cases of very large {beta}, while they agree with the results obtained with the old implementation of FLOW in lower-{beta} equilibria.
Energy Technology Data Exchange (ETDEWEB)
Hayami, H.; Kawaguchi, N. (Kyushu University, Fukuoka (Japan). Institute of Advanced Material Study); Sawae, M. (Nippon Mining Co. Ltd., Tokyo (Japan)); Nakamura, T. (Toshiba Corp., Tokyo (Japan))
1992-06-25
In this study, a low-solidity circular cascade, conformally transformed from a high-stagger linear cascade of double-circular-arc vanes with solidity of 0.69, was examined as a part of the diffuser system of a transonic centrifugal compressor. The blade loading of the cascade was also investigated by means of pressure measurement around the vane. Experiments were conducted by the testing apparatus of closed loop type compressor using fleon 12 under the condition of four rotating speeds between 15,000 and 19,000 r.p.m. Cascades with stagger angles of 69{degree} and 72{degree} were used. Consequently, it was found that the experimental data for the lift-coefficient of the vane were almost on a single straight line when plotted against angle-of-attack for a wide range of Mach numbers and flow angles. The maximum lift-coefficient of about 1.5 was recorded. It was also found that the vane functioned well even near the surge condition of the compressor. 6 refs., 6 figs.
Flegel, Ashlie B.
2014-01-01
The purpose of this thesis is to document the impact of incidence angle and Reynolds number variations on the three-dimensional flow field and midspan loss and turning of a two-dimensional section of a variable-speed power-turbine (VSPT) rotor blade. Aerodynamic measurements were obtained in a transonic linear cascade at NASA Glenn Research Center in Cleveland, Ohio. Steady-state data were obtained for 10 incidence angles ranging from +15.8deg to -51.0deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×105 to 2.12×106. Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan tota lpressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial-chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8deg and -36.7deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures.
Wang, Kon-Sheng Charles
1994-01-01
The design and development of an airborne flight-test experiment to study nonreacting gas jets injected transversely into transonic and supersonic crossflows is presented. Free-stream/crossflow Mach numbers range from 0.8 to 2.0. Planar laser-induced fluorescence (PLIF) of an iodine-seeded nitrogen jet is used to visualize the jet flow. Time-dependent images are obtained with a high-speed intensified video camera synchronized to the laser pulse rate. The entire experimental assembly is configured compactly inside a unique flight-test-fixture (FTF) mounted under the fuselage of the F-104G research aircraft, which serves as a 'flying wind tunnel' at NASA Dryden Flight Research Center. The aircraft is flown at predetermined speeds and altitudes to permit a perfectly expanded (or slightly underexpanded) gas jet to form just outside the FTF at each free-stream Mach number. Recorded gas jet images are then digitized to allow analysis of jet trajectory, spreading, and mixing characteristics. Comparisons will be made with analytical and numerical predictions. This study shows the viability of applying highly sophisticated groundbased flow diagnostic techniques to flight-test vehicle platforms that can achieve a wide range of thermo/fluid dynamic conditions. Realistic flow environments, high enthalpies, unconstrained flowfields, and moderate operating costs are also realized, in contrast to traditional wind-tunnel testing.
Reynolds, Robert M; Samonds, Robert I; Walker, John H
1957-01-01
An investigation has been made to determine the aerodynamic characteristics of the NACA 4-(5)(05)-041 four-blade, single-relation propeller and the NACA 4-(5)(05)-037 six- and eight-blade, dual-rotation propellers in combination with various spinners and NACA d-type spinner-cowling combinations at Mach numbers up to 0.84. Propeller force characteristics, local velocity distributions in the propeller planes, inlet pressure recoveries, and static-pressure distributions on the cowling surfaces were measured for a wide range of blade angles, advance ratios, and inlet-velocity ratios. Included are data showing: (a) the effect of extended cylindrical spinners on the characteristics of the single-rotation propeller, (b) the effect of variation of the difference in blade angle setting between the front and rear components of the dual-rotation propellers, (c) the negative- and static-thrust characteristics of the propellers with 1 series spinners, and (d) the effects of ideal- and platform-type propeller-spinner junctures on the pressure-recovery characteristics of the single-rotation propeller-spinner-cowling combination.
Pendergraft, O. C., Jr.; Schmeer, J. W.
1972-01-01
Twin-jet afterbody models were investigated by using two balances to measure the thrust-minus-total drag and the afterbody drag, separately, at static conditions and at Mach numbers up to 2.2 for an angle of attack of 0 deg. Hinged-flap convergent-divergent nozzles were tested at subsonic-cruise- and maximum-afterburning-power settings with a high-pressure air system used to provide jet-total-pressure ratios up to 20. Two nozzle lateral spacings were studied, using afterbodies with similar interfairing shapes but with different longitudinal cross-sectional area distributions. Alternate, blunter, interfairings with different shapes for the two spacings, which produced afterbodies having identical cross-sectional area progressions corresponding to an axisymmetric minimum wave-drag configuration, were also tested. The results indicate that the wide-spaced configurations improved the flow field around the nozzles, thereby reducing drag on the cruise nozzles; however, the increased surface and projected cross-sectional areas caused an increase in afterbody drag. Except for a slight advantage with cruise nozzles at subsonic speeds, the wide-spaced configurations had the higher total drag at all other test conditions.
Aeroacoustic computation of low mach number flow
Energy Technology Data Exchange (ETDEWEB)
Skriver Dahl, K. [Risoe National Laboratory, Roskilde (Denmark)
1997-12-31
The possibilities of applying a recently developed numerical technique to predict aerodynamically generated sound from wind turbines is explored. The technique is a perturbation technique that has the advantage that the underlying flow field and the sound field are computed separately. Solution of the incompressible, time dependent flow field yields a hydrodynamic density correction to the incompressible constant density. The sound field is calculated from a set of equations governing the inviscid perturbations about the corrected flow field. Here, the emphasis is placed on the computation of the sound field. The nonlinear partial differential equations governing the sound fields are solved numerically using an explicit MacCormack scheme. Two types of non-reflecting boundary conditions are applied; one based on the asymptotic solution of the governing equations and the other based on a characteristic analysis of the governing equations. The former condition is easy to use and it performs slightly better than the charcteristic based condition. The technique is applied to the problems of the sound generation of a co-rotating vortex pair, which is a quadrupole, and the viscous flow over a circular cylinder, which is a dipole. Numerical results agree very well with the analytical solution for the problem of the co-rotating vortex pair. Numerical results for the viscous flow over a cylinder are presented and evaluated qualitatively. (au)
Semidirect computations for transonic flow
Swisshelm, J. M.; Adamczyk, J. J.
1983-01-01
A semidirect method, driven by a Poisson solver, was developed for inviscid transonic flow computations. It is an extension of a recently introduced algorithm for solving subsonic rotational flows. Shocks are captured by implementing a form of artificial compressibility. Nonisentropic cases are computed using a shock tracking procedure coupled with the Rankine-Hugoniot relationships. Results are presented for both subsonic and transonic flows. For the test geometry, an unstaggered cascade of 20 percent thick circular arc airfoils at zero angle of attack, shocks are crisply resolved in supercritical situations and the algorithm converges rapidly. In addition, the convergence rate appears to be nearly independent of the entropy and vorticity production at the shock.
Energy Technology Data Exchange (ETDEWEB)
Handa, T.; Miyazato, Y.; Maasuda, M.; Matsuo, K. [Kyushu Univaersity, Fukuoka (Japan)
2000-05-25
Concerning the self-excited shock oscillation in a transonic diffuser flow, the response of a normal shock wave to a 'white' pressure disturbance, that is, a small pressure disturbance having the same power spectral density over every frequency, is analyzed by solving the equation representing shock displacement due to the disturbance. According to the present analysis, the spectral density distribution and peak frequency of the shock oscillation depend on Mach number just upstream of the shock wave and a non-dimensional parameter determined from the diffuser geometry at the time-mean shock location. The shock displacement power spectral density distributions obtained by the analysis agree very well with those measured in our experiment. A non-dimensional relationship between the peak frequency of the shock oscillation and diffuser geometry at the shock location, Eq. (18), is obtained, and this equation agrees well with the present and previous experimental results. (author)
Analysis of Unsteady Tip and Endwall Heat Transfer in a Highly Loaded Transonic Turbine Stage
Shyam, Vikram; Ameri, Ali; Chen, Jen-Ping
2010-01-01
In a previous study, vane-rotor shock interactions and heat transfer on the rotor blade of a highly loaded transonic turbine stage were simulated. The geometry consists of a high pressure turbine vane and downstream rotor blade. This study focuses on the physics of flow and heat transfer in the rotor tip, casing and hub regions. The simulation was performed using the Unsteady Reynolds-Averaged Navier-Stokes (URANS) code MSU-TURBO. A low Reynolds number k-epsilon model was utilized to model turbulence. The rotor blade in question has a tip gap height of 2.1 percent of the blade height. The Reynolds number of the flow is approximately 3x10(exp 6) per meter. Unsteadiness was observed at the tip surface that results in intermittent "hot spots". It is demonstrated that unsteadiness in the tip gap is governed by inviscid effects due to high speed flow and is not strongly dependent on pressure ratio across the tip gap contrary to published observations that have primarily dealt with subsonic tip flows. The high relative Mach numbers in the tip gap lead to a choking of the leakage flow that translates to a relative attenuation of losses at higher loading. The efficacy of new tip geometry is discussed to minimize heat flux at the tip while maintaining choked conditions. In addition, an explanation is provided that shows the mechanism behind the rise in stagnation temperature on the casing to values above the absolute total temperature at the inlet. It is concluded that even in steady mode, work transfer to the near tip fluid occurs due to relative shearing by the casing. This is believed to be the first such explanation of the work transfer phenomenon in the open literature. The difference in pattern between steady and time-averaged heat flux at the hub is also explained.
Phillips, W. P.; Fournier, R. H.
1985-01-01
Wind-tunnel tests were conducted at Mach 1.5 to 2.5 to determine the effect of modifications designed to extend the forward center-of-gravity trim capability on the static longitudal and lateral directional characteristics of a Space shuttle 140 A/B orbiter model (0.01 scale). The modifications consisted of a forward-extended wing fillet, a flat plate canard, and a blended canard. The investigation was conducted in the low Mach number test section of the Langley unitary plan wind tunnel at a Reynolds number of approximately 2.15 million based on the fuselage reference length. The test angle of attack range was -1 deg to 32 deg and the sideslip angles were 0 deg and 5 deg.
Realizability of stationary spherically symmetric transonic accretion
Ray, A K; Ray, Arnab K.
2002-01-01
The spherically symmetric stationary transonic (Bondi) flow is considered a classic example of an accretion flow. This flow, however, is along a separatrix, which is usually not physically realizable. We demonstrate, using a pedagogical example, that it is the dynamics which selects the transonic flow.
Flegel-McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.
2013-01-01
Aerodynamic measurements obtained in a transonic linear cascade were used to assess the impact of large incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Steady-state data were obtained for ten incidence angles ranging from +15.8 deg to -51.0 deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×10(exp 5) to 2.12×10(exp 6). Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8 deg and -36.7 deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures. Tests were conducted in the NASA Glenn Transonic Turbine Blade Cascade Facility. The measurements reflect strong secondary flows associated with the high aerodynamic loading levels at large positive incidence angles and an increase in loss levels with decreasing Reynolds number. The secondary flows decrease with negative incidence as the blade becomes unloaded. Transitional flow is admitted in this low inlet turbulence dataset, making it a challenging CFD test case. The dataset will be used to advance understanding of the aerodynamic
Miller, Rolf W.; Argrow, Brian M.; Center, Kenneth B.; Brauckmann, Gregory J.; Rhode, Matthew N.
1998-01-01
The NASA Langley Research Center Unitary Plan Wind Tunnel and the 20-Inch Mach 6 Tunnel were used to test two osculating cones waverider models. The Mach-4 and Mach-6 shapes were generated using the interactive design tool WIPAR. WIPAR performance predictions are compared to the experimental results. Vapor screen results for the Mach-4 model at the on- design Mach number provide visual verification that the shock is attached along the entire leading edge, within the limits of observation. WIPAR predictions of pressure distributions and aerodynamic coefficients show general agreement with the corresponding experimental values.
Brown, C. A., Jr.; Campbell, J. F.
1973-01-01
An investigation was conducted to obtain flow properties in the wake of a preliminary configuration of the Viking '75 Entry Vehicle at Mach numbers from 0.20 to 1.20 and at angles of attack of 0 deg, 5 deg, and 10 deg. The wake flow properties were calculated from total and static pressures measured with a pressure rake at longitudinal stations varying from 1.50 to 11.00 body diameters, and are presented in tabulated and plotted form. The wake properties were essentially symmetrical about the X-axis at alpha = 0 deg and the profiles were shifted away from the X-axis at angles of attack. An unexpected reduction in wake property ratios occurred as the Mach number increased from 0.60 to 1.00; these ratios then increased as the Mach number increased to 1.20. The reduction was present for all the longitudinal stations of the tests and decreased with increased longitudinal distance.
Runckel, Jack F.; Hieser, Gerald
1961-01-01
An investigation has been conducted at the Langley 16-foot transonic tunnel to determine the loading characteristics of flap-type ailerons located at inboard, midspan, and outboard positions on a 45 deg. sweptback-wing-body combination. Aileron normal-force and hinge-moment data have been obtained at Mach numbers from 0.80 t o 1.03, at angles of attack up to about 27 deg., and at aileron deflections between approximately -15 deg. and 15 deg. Results of the investigation indicate that the loading over the ailerons was established by the wing-flow characteristics, and the loading shapes were irregular in the transonic speed range. The spanwise location of the aileron had little effect on the values of the slope of the curves of hinge-moment coefficient against aileron deflection, but the inboard aileron had the greatest value of the slope of the curves of hinge-moment coefficient against angle of attack and the outboard aileron had the least. Hinge-moment and aileron normal-force data taken with strain-gage instrumentation are compared with data obtained with pressure measurements.
Control of the NASA Langley 16-Foot Transonic Tunnel with the Self-Organizing Feature Map
Motter, Mark A.
1998-01-01
A predictive, multiple model control strategy is developed based on an ensemble of local linear models of the nonlinear system dynamics for a transonic wind tunnel. The local linear models are estimated directly from the weights of a Self Organizing Feature Map (SOFM). Local linear modeling of nonlinear autonomous systems with the SOFM is extended to a control framework where the modeled system is nonautonomous, driven by an exogenous input. This extension to a control framework is based on the consideration of a finite number of subregions in the control space. Multiple self organizing feature maps collectively model the global response of the wind tunnel to a finite set of representative prototype controls. These prototype controls partition the control space and incorporate experimental knowledge gained from decades of operation. Each SOFM models the combination of the tunnel with one of the representative controls, over the entire range of operation. The SOFM based linear models are used to predict the tunnel response to a larger family of control sequences which are clustered on the representative prototypes. The control sequence which corresponds to the prediction that best satisfies the requirements on the system output is applied as the external driving signal. Each SOFM provides a codebook representation of the tunnel dynamics corresponding to a prototype control. Different dynamic regimes are organized into topological neighborhoods where the adjacent entries in the codebook represent the minimization of a similarity metric which is the essence of the self organizing feature of the map. Thus, the SOFM is additionally employed to identify the local dynamical regime, and consequently implements a switching scheme than selects the best available model for the applied control. Experimental results of controlling the wind tunnel, with the proposed method, during operational runs where strict research requirements on the control of the Mach number were met, are
Schrenk, Markus
2011-01-01
In his Contributions to the Analysis of the Sensations (Mach 1885) the phenomenalist philosopher Ernst Mach confronts us with a difficulty: “If we regard the Ego as a real unity, we become involved in the following dilemma: either we must set over against the Ego a world of unknowable entities […] or we must regard the whole world, the Egos of other people included, as comprised in our own Ego.” (Mach 1885: 21) In other words, if we start from a phenomenalist viewpoint, i.e., if we believ...
Fourier time spectral method for subsonic and transonic flows
Institute of Scientific and Technical Information of China (English)
Lei Zhan; Feng Liu; Dimitri Papamoschou
2016-01-01
The time accuracy of the exponentially accu-rate Fourier time spectral method (TSM) is examined and compared with a conventional 2nd-order backward differ-ence formula (BDF) method for periodic unsteady flows. In particular, detailed error analysis based on numerical com-putations is performed on the accuracy of resolving the local pressure coefficient and global integrated force coefficients for smooth subsonic and non-smooth transonic flows with moving shock waves on a pitching airfoil. For smooth sub-sonic flows, the Fourier TSM method offers a significant accuracy advantage over the BDF method for the predic-tion of both the local pressure coefficient and integrated force coefficients. For transonic flows where the motion of the discontinuous shock wave contributes significant higher-order harmonic contents to the local pressure fluctuations, a sufficient number of modes must be included before the Fourier TSM provides an advantage over the BDF method. The Fourier TSM, however, still offers better accuracy than the BDF method for integrated force coefficients even for transonic flows. A problem of non-symmetric solutions for symmetric periodic flows due to the use of odd numbers of intervals is uncovered and analyzed. A frequency-searching method is proposed for problems where the frequency is not known a priori. The method is tested on the vortex shedding problem of the flow over a circular cylinder.
Isaev, Sergey; Baranov, Paul; Popov, Igor; Sudakov, Alexander; Usachov, Alexander
2017-03-01
The modified SST model (2005) is verified using Rodi- Leschziner-Isaev's approach and the multiblock computational technologies are validated in the VP2/3 code on different-structure overlapping grids by comparing the numerical predictions with the experimental data on transonic flow around an NACA0012 airfoil at an angle of attack of 4o for M=0.7 and Re=4×106. It is proved that the aerodynamic characteristics of a thick (20% of the chord) MQ airfoil mounted at an angle of attack of 2o for Re=107 and over the Mach number range 0.3-0.55 are significantly improved because an almost circular small-size (0.12) vortex cell with a defined volumetric flow rate coefficient of 0.007 during slot suction has been located on the upper airfoil section and an intense trapped vortex has been formed in it. A detailed analysis of buffeting within the self-oscillatory regime of flow around the MQ airfoil with a vortex cell has demonstrated the periodic changes in local and integral characteristics; the lift and the aerodynamic efficiency remain quite high, but inferior to the similar characteristics at M=0.55. It is found that the vortex cell at M=0.7 is inactive, and the aerodynamic characteristics of the MQ airfoil with a vortex cell are close to those of a smooth airfoil without a cell.
Transonic flutter study of a wind-tunnel model of a supercritical wing with/without winglet
Ruhlin, C. L.; Rauch, F. J., Jr.; Waters, C.
1982-01-01
The scaled flutter model was a 1/6.5-size, semispan version of a supercritical wing (SCW) proposed for an executive-jet-transport airplane. The model was tested cantilever-mounted with a normal wingtip, a wingtip with winglet, and a normal wingtip ballasted to simulate the winglet mass properties. Flutter and aerodynamic data were acquired at Mach numbers from 0.6 to 0.95. The measured transonic flutter speed boundary for each wingtip configuration had roughly the same shape with a minimum flutter speed near M = 0.82. The winglet addition and wingtip mass ballast decreased the wing flutter speed by about 7 and 5%, respectively; thus, the winglet effect on flutter was more a mass effect than an aerodynamic effect. Flutter characteristics calculated using a doublet-lattice analysis (which included interference effects) were in good agreement with the experimental results up to M = 0.82. Comparisons of measured static aerodynamic data with predicted data indicated that the model was aerodynamically representative of the airplane SCW.
Energy Technology Data Exchange (ETDEWEB)
Fortin, T
2006-05-15
This work deals with the discretization of Navier-Stokes equations using different finite element methods adapted to the problem of two-phase flows. These methods must be of high order to limit the presence of spurious flows (which contradict the establishment of a physical equilibrium) and to verify energy conservation properties. Several solutions are proposed which seem to fulfill these expectations. A reformulation of the six-equation system adapted to low Mach two-phase flows has been also proposed. These methods have been implemented into the Trio-U code of CEA Grenoble, but have been tested only on simple 'academic' configurations. (J.S.)
Erickson, Gary E.; Gonzalez, Hugo A.
2006-01-01
A pressure-sensitive paint (PSP) technique was applied in a wind tunnel experiment in the NASA Langley Research Center 8-Foot Transonic Pressure Tunnel to study the effect of wing fillets on the global vortex induced surface static pressure field about a sharp leading-edge 76 deg./40 deg. double delta wing, or strake-wing, model at subsonic and transonic speeds. Global calibrations of the PSP were obtained at M(sub infinity) = 0.50, 0.70, 0.85, 0.95, and 1.20, a Reynolds number per unit length of 2.0 million, and angles of attack from 10 degrees to 20 degrees using an insitu method featuring the simultaneous acquisition of electronically scanned pressures (ESP) at discrete locations on the model. The mean error in the PSP measurements relative to the ESP data was approximately 2 percent or less at M(sub infinity) = 0.50 to 0.85 but increased to several percent at M(sub infinity) =0.95 and 1.20. The PSP pressure distributions and pseudo-colored, planform-view pressure maps clearly revealed the vortex-induced pressure signatures at all Mach numbers and angles of attack. Small fillets having parabolic or diamond planforms situated at the strake-wing intersection were respectively designed to manipulate the vortical flows by removing the leading-edge discontinuity or introducing additional discontinuities. The fillets caused global changes in the vortex-dominated surface pressure field that were effectively captured in the PSP measurements. The vortex surface pressure signatures were compared to available off-surface vortex cross-flow structures obtained using a laser vapor screen (LVS) flow visualization technique. The fillet effects on the PSP pressure distributions and the observed leading-edge vortex flow characteristics were consistent with the trends in the measured lift, drag, and pitching moment coefficients.
Rivers, Melissa; Quest, Juergen; Rudnik, Ralf
2015-01-01
Experimental aerodynamic investigations of the NASA Common Research Model have been conducted in the NASA Langley National Transonic Facility, the NASA Ames 11-ft wind tunnel, and the European Transonic Wind Tunnel. In the NASA Ames 11-ft wind tunnel, data have been obtained at only a chord Reynolds number of 5 million for a wing/body/tail = 0 degree incidence configuration. Data have been obtained at chord Reynolds numbers of 5, 19.8 and 30 million for the same configuration in the National Transonic Facility and in the European Transonic Facility. Force and moment, surface pressure, wing bending and twist, and surface flow visualization data were obtained in all three facilities but only the force and moment and surface pressure data are presented herein.
Jernell, Lloyd S.
1961-01-01
An investigation w a s made i n the Langley Unitary Plan wind tunnel o determine the effects of fin area and the effects of antennas and w iring tunnels on the static longitudinal and lateral stability of a 0 .10- scale model of a three- stage configuration of the Scout vehicle. The tests were performed at Mach numbers of 2.29, 2.96, 3.96, and 4. 65 6 and at Reynolds numbers of about 3.5 X 10 per foot.
Slow light Mach-Zehnder fiber interferometer
Institute of Scientific and Technical Information of China (English)
Yundong Zhang; Jinfang Wang; Xuenan Zhang; Hao Wu; Yuanxue Cai; Jing Zhang; Ping Yuan
2012-01-01
A slow light structure Mach-Zehnder fiber interferometer is theoretically demonstrated.The sensitivity of the interferometer is significantly enhanced by the dispersion of the slow light structure.The numerical results show that the sensitivity enhancement factor varies with the coupling coefficient and reaches its maximum under critical coupling conditions.Interferometers have been investigated in relation to their applications in fields such as metrology[1],optical sensing[2],optical communication[3,4],quantum information processing[5],and biomedical engineering[6].A number of schemes have been proposed to improve the performance of interferometers[7],such as using photonic crystal structures to minimize the size of on-chip devices[8],utilizing the dispersive property of semiconductor to enhance the spectral sensitivity of interferometers[9,10],utilizing slow light medium to enhance the resolution of Fourier transform interferometer[11],exploiting fast light medium or slow light structure to increase the rotation sensitivity of a Sagnac interferometer[12,13],enhancing the transmittance of the Mach-Zehnder interferometer (MZI) in the slow light region by gratings[14],and using liquid crystal light valve to derive high sensitivity interferometers[15].%A slow light structure Mach-Zehnder fiber interferometer is theoretically demonstrated. The sensitivity of the interferometer is significantly enhanced by the dispersion of the slow light structure. The numerical results show that the sensitivity enhancement factor varies with the coupling coefficient and reaches its maximum under critical coupling conditions.
Energy Technology Data Exchange (ETDEWEB)
Kegalj, Martin
2013-11-01
In axial turbines tip leakage forms a large portion of the overall losses. Applying a shroud is very aerodynamically useful, but the higher mechanical loads of the revolving rotor blading exposed to a high thermal load and the higher costs suggest a shroudless configuration is better. The main parameter in the tip leakage loss is the tip gap height, which cannot be reduced arbitrarily as a running gap is necessary due to thermal expansion and vibration of the jet engine. The pressure ratio between pressure and suction of the rotor blade forces the fluid over the blade tip and leads to the formation of the tip leakage vortex. Reduced turning and losses caused by vortices and subsequent mixing are responsible for the reduced efficiency. Using a squealer cavity on the flat blade tip is a feasible way to reduce the aerodynamic losses. A portion of the kinetic energy of the tip leakage flow is dissipated while entering the cavity; the flow exiting the cavity enters the passage with reduced momentum and reduced tip gap mass flow. A 1(1)/(2) stage low mach number turbine was used to investigate the influence of tip geometry. Aerodynamic measurements, performed with five-hole probes, two-component hot-wire anemometer, unsteady wall pressure sensors, stereo and borescopic particle-image-velocimetry setups and oil and dye flow visualization, found small differences in the flow velocities and angles between the flat and squealer tip configuration in the measurement planes downstream of the rotor. The measurement uncertainty proves the difficulty of determining the influence of the squealer cavity on the blade row outflow with global measurement data. To gather information on the flow close to the casing inside the rotor passage is only possible with non-intrusive laser measurement techniques. Comparison of the different tip geometries is still difficult due to the small differences in the absolute flow data. The use of the {lambda}{sub 2} vortex criterion enables an objective
Computer program calculates transonic velocities in turbomachines
Katsanis, T.
1971-01-01
Computer program, TSONIC, combines velocity gradient and finite difference methods to obtain numerical solution for ideal, transonic, compressible flow for axial, radial, or mixed flow cascade of turbomachinery blades.
A model for transonic plasma flow
Energy Technology Data Exchange (ETDEWEB)
Guazzotto, Luca, E-mail: luca.guazzotto@rochester.edu [Laboratory for Laser Energetics, University of Rochester, Rochester, New York 14623 (United States); Hameiri, Eliezer, E-mail: hameiri@cims.nyu.edu [Courant Institute of Mathematical Sciences, New York University, New York, New York 10012 (United States)
2014-02-15
A linear, two-dimensional model of a transonic plasma flow in equilibrium is constructed and given an explicit solution in the form of a complex Laplace integral. The solution indicates that the transonic state can be solved as an elliptic boundary value problem, as is done in the numerical code FLOW [Guazzotto et al., Phys. Plasmas 11, 604 (2004)]. Moreover, the presence of a hyperbolic region does not necessarily imply the presence of a discontinuity or any other singularity of the solution.
Effects of Porous Throat on Transonic Diffuser
屋我, 実; 永井, 實; 富田, 教夫; 芳賀, 剛; 宮良, 透; Yaga, Minoru; Nagai, Minoru; Tomita, Norio; Haga, Tsuyoshi; Miyara, Tooru
1995-01-01
The effects of the porous throat on a transonic diffuser were investigated experimentally by wall static pressure measurements and by schlieren optical observations. The porous throat consists of a wall with 126 holes and a cavity underneath it so that the flow around the shock wave can circulate through the porous wall. The results show that no shock wave was observed at 80% of the porous region from the throat and that the pressure fluctuations in the transonic diffuser were greatly reduced...
Transonic analysis of canted winglets
Rosen, B. S.
1984-01-01
A computational method developed to provide a transonic analysis for upper/lower surface wing-tip mounted winglets is described. Winglets with arbitrary planform, cant and toe angle, and airfoil section can be modeled. The embedded grid approach provides high flow field resolution and the required geometric flexibility. In particular, coupled Cartesian/cylindrical grid systems are used to model the complex geometry presented by canted upper/lower surface winglets. A new rotated difference scheme is introduced in order to maintain the stability of the small-disturbance formulation in the presence of large spanwise velocities. Wing and winglet viscous effects are modeled using a two-dimensional 'strip' boundary layer analysis. Correlations with wind tunnel and flight test data for three transport configurations are included.
Transonic Drag Reduction Through Trailing-Edge Blowing on the FAST-MAC Circulation Control Model
Chan, David T.; Jones, Gregory S.; Milholen, William E., II; Goodliff, Scott L.
2017-01-01
A third wind tunnel test of the FAST-MAC circulation control semi-span model was completed in the National Transonic Facility at the NASA Langley Research Center where the model was configured for transonic testing of the cruise configuration with 0deg flap detection to determine the potential for transonic drag reduction with the circulation control blowing. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged ap. Recent upgrades to transonic semi-span flow control testing at the NTF have demonstrated an improvement to overall data repeatability, particularly for the drag measurement, that allows for increased confidence in the data results. The static thrust generated by the blowing slot was removed from the wind-on data using force and moment balance data from wind-o thrust tares. This paper discusses the impact of the trailing-edge blowing to the transonic aerodynamics of the FAST-MAC model in the cruise configuration, where at flight Reynolds numbers, the thrust-removed corrected data showed that an overall drag reduction and increased aerodynamic efficiency was realized as a consequence of the blowing.
Ball, J. W.; Lindahl, R. H.
1976-01-01
The purpose of the test was to investigate the nature of the Orbiter boundary layer characteristics at angles of attack from -4 to 32 degrees at a Mach number of 4.6. The effect of large grit, employed as transition strips, on both the nature of the boundary layer and the force and moment characteristics were investigated along with the effects of large negative elevon deflection on lee side separation. In addition, laminar and turbulent boundary layer separation phenomena which could cause asymmetric flow separation were investigated.
Transonic behaviour and stability analysis of quasi-viscous black hole accretion
Ananda, Deepika B; Das, Tapas K; Maity, Ishita; Nag, Sankhasubhra
2015-01-01
Analytical studies of black hole accretion usually presumes the stability of the stationary transonic configuration. Various authors in the past several decades demonstrated the validity of such an assumption for inviscid hydrodynamic flow. Inviscid approximation is a reasonable approach for low angular momentum advection dominated flow in connection to certain supermassive black holes at the centres of the galaxies (including our own) fed from a number of stellar donors. Introduction of a weak viscosity, as a first order linear correction involving the viscosity parameter, however, may sometimes provide a more detail understanding of the observed black hole spectra. The transonic behaviour of the stationary solutions have been studied for the aforementioned quasi-viscous accretion for all possible geometric configurations of axisymmetric flow. For a sufficiently low range of the viscosity parameter, transonic solutions containing one or three critical points have been found for allowed ranges in the astrophy...
Wind tunnel wall interference effects on a supercritical airfoil at transonic speeds
Blackwell, J. A., Jr.; Pounds, G. A.
1976-01-01
Wind tunnel tests of a 10% supercritical airfoil have been conducted in the Lockheed Compressible Flow Facility at transonic speeds to determine the effects of varying wind tunnel wall porosity on airfoil performance. Wall configurations ranging in porosity from 1.3% to 10% were investigated at Reynolds numbers of 7 to 30 million. Experimental data presented to show the effect of varying wall porosity include airfoil surface pressures, airfoil forces, and wind tunnel wall pressures. Utilizing the experimental results, an assessment of the applicability of current subcritical theoretical methods to predict wall interference corrections in subsonic and transonic flows is made.
A Bibliography of Transonic Dynamics Tunnel (TDT) Publications
Doggett, Robert V.
2016-01-01
The Transonic Dynamics Tunnel (TDT) at the National Aeronautics and Space Administration's (NASA) Langley Research Center began research operations in early 1960. Since that time, over 600 tests have been conducted, primarily in the discipline of aeroelasticity. This paper presents a bibliography of the publications that contain data from these tests along with other reports that describe the facility, its capabilities, testing techniques, and associated research equipment. The bibliography is divided by subject matter into a number of categories. An index by author's last name is provided.
Unsteady transonic flow over cascade blades
Surampudi, S. P.; Adamczyk, J. J.
1986-01-01
An attempt is made to develop an efficient staggered cascade blade unsteady aerodynamics model for the neighborhood of March 1, representing the blade row by a rectilinear two-dimensional cascade of thin, flat plate airfoils. The equations of motion are derived on the basis of linearized transonic small perturbation theory, and an analytical solution is obtained by means of the Wiener-Hopf procedure. Making use of the transonic similarity law, the results obtained are compared with those of other linearized cascade analyses. A parametric study is conducted to find the effects of reduced frequency, stagger angle, solidity, and the location of the pitching axis on cascade stability.
Phillips, W. P.; Fournier, R. H.
1979-01-01
Supersonic aerodynamic characteristics are presented for the 140A/B space shuttle orbiter configuration (0.010 scale) and for the configuration modified to incorporate geometry changes in the wing planform fillet region. The modifications designed to extend the orbiter's longitudinal trim capability to more forward center-of-gravity locations, included reshaping of the baseline wing planform fillet and adding canards. The investigation was made in the high Mach number test section of the Langley Unitary Plan Wind Tunnel at a Reynolds number of approximately 2.2 million based on fuselage reference length. The angle-of-attack range for the investigation extended from -1 deg to 31 deg. Data were obtained with the elevators and body flap deflected at appropriate negative and positive conditions to assess the trim limits.
National Aeronautics and Space Administration — The need to accurately predict aeroelastic phenomenon for a wide range of Mach numbers is a critical step in the design process of any aerospace vehicle. Complex...
ANALYSIS OF TRANSONIC FLOW PAST CUSPED AIRFOILS
Directory of Open Access Journals (Sweden)
Jiří Stodůlka
2015-06-01
Full Text Available Transonic flow past two cusped airfoils is numerically solved and achieved results are analyzed by means of flow behavior and oblique shocks formation.Regions around sharp trailing edges are studied in detail and parameters of shock waves are solved and compared using classical shock polar approach and verified by reduction parameters for symmetric configurations.
Mach-like capillary-gravity wakes.
Moisy, Frédéric; Rabaud, Marc
2014-08-01
We determine experimentally the angle α of maximum wave amplitude in the far-field wake behind a vertical surface-piercing cylinder translated at constant velocity U for Bond numbers Bo(D)=D/λ(c) ranging between 0.1 and 4.2, where D is the cylinder diameter and λ(c) the capillary length. In all cases the wake angle is found to follow a Mach-like law at large velocity, α∼U(-1), but with different prefactors depending on the value of Bo(D). For small Bo(D) (large capillary effects), the wake angle approximately follows the law α≃c(g,min)/U, where c(g,min) is the minimum group velocity of capillary-gravity waves. For larger Bo(D) (weak capillary effects), we recover a law α∼√[gD]/U similar to that found for ship wakes at large velocity [Rabaud and Moisy, Phys. Rev. Lett. 110, 214503 (2013)]. Using the general property of dispersive waves that the characteristic wavelength of the wave packet emitted by a disturbance is of order of the disturbance size, we propose a simple model that describes the transition between these two Mach-like regimes as the Bond number is varied. We show that the new capillary law α≃c(g,min)/U originates from the presence of a capillary cusp angle (distinct from the usual gravity cusp angle), along which the energy radiated by the disturbance accumulates for Bond numbers of order of unity. This model, complemented by numerical simulations of the surface elevation induced by a moving Gaussian pressure disturbance, is in qualitative agreement with experimental measurements.
Sandford, M. C.; Ricketts, R. H.; Cazier, F. W., Jr.
1980-01-01
A supercritical wing with an aspect ratio of 10.76 and with two trailing-edge oscillating control surfaces is described. The semispan wing is instrumented with 252 static orifices and 164 in situ dynamic-pressure gages for studying the effects of control-surface position and motion on steady- and unsteady-pressures at transonic speeds. Results from initial tests conducted in the Langley Transonic Dynamics Tunnel at two Reynolds numbers are presented in tabular form.
Directory of Open Access Journals (Sweden)
Haojie Liu
2016-01-01
Full Text Available The paper presents a digital adaptive controller of recurrent neural networks for the active flutter suppression of a wing structure over a wide transonic range. The basic idea behind the controller is as follows. At first, the parameters of recurrent neural networks, such as the number of neurons and the learning rate, are initially determined so as to suppress the flutter under a specific flight condition in the transonic regime. Then, the controller automatically adjusts itself for a new flight condition by updating the synaptic weights of networks online via the real-time recurrent learning algorithm. Hence, the controller is able to suppress the aeroelastic instability of the wing structure over a range of flight conditions in the transonic regime. To demonstrate the effectiveness and robustness of the controller, the aeroservoelastic model of a typical fighter wing with a tip missile was established and a single-input/single-output controller was synthesized. Numerical simulations of the open/closed-loop aeroservoelastic simulations were made to demonstrate the efficacy of the adaptive controller with respect to the change of flight parameters in the transonic regime.
Uncertainty Quantification of Turbulence Model Closure Coefficients for Transonic Wall-Bounded Flows
Schaefer, John; West, Thomas; Hosder, Serhat; Rumsey, Christopher; Carlson, Jan-Renee; Kleb, William
2015-01-01
The goal of this work was to quantify the uncertainty and sensitivity of commonly used turbulence models in Reynolds-Averaged Navier-Stokes codes due to uncertainty in the values of closure coefficients for transonic, wall-bounded flows and to rank the contribution of each coefficient to uncertainty in various output flow quantities of interest. Specifically, uncertainty quantification of turbulence model closure coefficients was performed for transonic flow over an axisymmetric bump at zero degrees angle of attack and the RAE 2822 transonic airfoil at a lift coefficient of 0.744. Three turbulence models were considered: the Spalart-Allmaras Model, Wilcox (2006) k-w Model, and the Menter Shear-Stress Trans- port Model. The FUN3D code developed by NASA Langley Research Center was used as the flow solver. The uncertainty quantification analysis employed stochastic expansions based on non-intrusive polynomial chaos as an efficient means of uncertainty propagation. Several integrated and point-quantities are considered as uncertain outputs for both CFD problems. All closure coefficients were treated as epistemic uncertain variables represented with intervals. Sobol indices were used to rank the relative contributions of each closure coefficient to the total uncertainty in the output quantities of interest. This study identified a number of closure coefficients for each turbulence model for which more information will reduce the amount of uncertainty in the output significantly for transonic, wall-bounded flows.
Effect of cavity on shock oscillation in transonic flow over RAE2822 supercritical airfoil
Rahman, M. Rizwanur; Labib, Md. Itmam; Hasan, A. B. M. Toufique; Ali, M.; Mitsutake, Y.; Setoguchi, T.
2016-07-01
Transonic flow past a supercritical airfoil is strongly influenced by the interaction of shock wave with boundary layer. This interaction induces unsteady self-sustaining shock wave oscillation, flow instability, drag rise and buffet onset which limit the flight envelop. In the present study, a computational analysis has been carried out to investigate the flow past a supercritical RAE2822 airfoil in transonic speeds. To control the shock wave oscillation, a cavity is introduced on the airfoil surface where shock wave oscillates. Different geometric configurations have been investigated for finding optimum cavity geometry and dimension. Unsteady Reynolds averaged Navier-Stokes equations (RANS) are computed at Mach 0.729 with an angle of attack of 5°. Computed results are well validated with the available experimental data in case of baseline airfoil. However, in case of airfoil with control cavity; it has been observed that the introduction of cavity completely suppresses the unsteady shock wave oscillation. Further, significant drag reduction and successive improvement of aerodynamic performance have been observed in airfoil with shock control cavity.
Energy Technology Data Exchange (ETDEWEB)
Stetter, H.; Urban, B.; Bauer, H.
1999-12-01
For experimental investigations on shock-induced flutter in a linear transonic turbine cascade an elastic suspension system has been developed so that only aerodynamic coupling occurs in the system. The test facility uses super-heated steam as working fluid and enables Mach and Reynolds numbers to vary independently. The investigated cascade consists of seven prismatic blades. The profiles are taken from the tip section of a transonic low pressure steam turbine blade. Each blade is suspended by an elastic spring system which allows the respective blade to vibrate in a mode similar to the real blade's first bending mode. The examination mainly deals with the oscillatory behavior of the blades with respect to a variation in the isentropic outlet Mach number. In addition, the complex shock-boundary-layer interactions on the blades' suction sides are described. Flow computations are run by a finite volume Navier-Stokes solver that accounts for moving boundaries. A volume grid generator is integrated into the flow solver producing combined O- and H-type grids. Turbulence is modeled by a k-{epsilon} turbulence model using wall functions because of performance reasons. Some acceleration techniques for unsteady flow computations are investigated. Shock oscillations which occur on a DCA profile are simulated. For the simulation of the experimental setup the blade motions are prescribed. (orig.) [German] Fuer ein Spitzenschnittprofil einer ND-Endstufenschaufel sowie eine Gasturbinenschaufel wurden lineare Gitter bestehend aus jeweils sieben elastisch aufgehaengten Schaufeln entwickelt. Die Konstruktion der Aufhaengung gewaehrleistet, dass die Schaufeln lediglich durch aerodynamische Effekte zu freien Schwingungen angeregt werden koennen und in ihrem Schwingungsverhalten der ersten Biegeschwingung ihrer Originalschaufeln entsprechen. Ein gestimmtes und ungestimmtes ebenes Gitter wurde in einem Dampfversuchsstand mit Heissdampf durchstroemt, wobei das
Energy Technology Data Exchange (ETDEWEB)
Stetter, H.; Urban, B.; Bauer, H.
1999-12-01
For experimental investigations on shock-induced flutter in a linear transonic turbine cascade an elastic suspension system has been developed so that only aerodynamic coupling occurs in the system. The test facility uses super-heated steam as working fluid and enables Mach and Reynolds numbers to vary independently. The investigated cascade consists of seven prismatic blades. The profiles are taken from the tip section of a transonic low pressure steam turbine blade. Each blade is suspended by an elastic spring system which allows the respective blade to vibrate in a mode similar to the real blade's first bending mode. The examination mainly deals with the oscillatory behavior of the blades with respect to a variation in the isentropic outlet Mach number. In addition, the complex shock-boundary-layer interactions on the blades' suction sides are described. Flow computations are run by a finite volume Navier-Stokes solver that accounts for moving boundaries. A volume grid generator is integrated into the flow solver producing combined O- and H-type grids. Turbulence is modeled by a k-{epsilon} turbulence model using wall functions because of performance reasons. Some acceleration techniques for unsteady flow computations are investigated. Shock oscillations which occur on a DCA profile are simulated. For the simulation of the experimental setup the blade motions are prescribed. (orig.) [German] Fuer ein Spitzenschnittprofil einer ND-Endstufenschaufel sowie eine Gasturbinenschaufel wurden lineare Gitter bestehend aus jeweils sieben elastisch aufgehaengten Schaufeln entwickelt. Die Konstruktion der Aufhaengung gewaehrleistet, dass die Schaufeln lediglich durch aerodynamische Effekte zu freien Schwingungen angeregt werden koennen und in ihrem Schwingungsverhalten der ersten Biegeschwingung ihrer Originalschaufeln entsprechen. Ein gestimmtes und ungestimmtes ebenes Gitter wurde in einem Dampfversuchsstand mit Heissdampf durchstroemt, wobei das
跨声速风洞中的超临界翼型速度场测量%Velocity Field Measurement of Supercritical Airfoil in Transonic Wind Tunnel
Institute of Scientific and Technical Information of China (English)
王猛; 李玉军; 衷洪杰
2015-01-01
针对跨声速风洞实验环境，发展适用于高速流动测量的粒子图像测速（PIV）实验技术，从而实现对超临界翼型的非定常流动进行准确的速度场测量。对跨声速风洞中的PIV测量系统进行了如下改进工作：改进粒子播撒装置；对片光光路布局进行了优化；对模型表面进行了防漫反射处理；提高图像位移场计算精度。使用RAE 2822超临界翼型，在FL-1风洞进行实验，来流马赫数为0.6～0.92，雷诺数为1.86×106～2.58×106。结果表明，利用PIV测得的速度场能够有效捕捉激波位置，并且与表面平均压力分布一致，通过对PIV数据进行统计分析可得到激波振荡的类型及幅度等特征。%A PIV measurement system had been developed to measure the unsteady lfow structure of the supercritical airfoil in transonic wind tunnel. The following improvements had been made to the PIV measurement system for transonic wind tunnel: improved particle seeding device, optimized the light path arrangement, anti-diffuse relfection treatment on the model surface, and increased the image displacement ifeld calculation accuracy. An experiment was conducted at FL-1 wind tunnel to investigate the behavior of transient shock wave/boundary layer interaction on a supercritical airfoil (RAE 2822). The Mach number was between 0.6 and 0.92, and Reynolds number between 1.86×106 and 2.58×106. The results show that the velocity ifeld measured by PIV can effectively capture the shock position, and the average pressure distribution is consistent with that of the PIV data. The type and magnitude of the shock wave can be obtained by statistical analysis.
Complex configuration analysis at transonic speeds
Boppe, C. W.; Aidala, P. V.
1980-01-01
Advanced performance requirements of new combat and transport aircraft together with design time constraints intensify the development and application of three dimensional computational analyses. A computational method which was developed for the specific purpose of providing an engineering analysis of complex aircraft configurations at transonic speeds. Particular attention is given to the recently incorporated wing viscous interaction and canard capabilities. The treatment of fuselage fairings, nacelles, and pylons is reviewed. The means for keeping computing resources at reasonable levels are identified. Three configurations were selected for correlations with experimental data. Taken together, the comparisons illustrate the full extent of current analysis capabilities. The configurations include: (1) a wing fuselage canard fighter; (2) a transport with fuselage fairings, four nacelles, four pylons; and (3) a space vehicle which includes an external fuel tank and rocket boosters (transonic launch configuration).
Analysis of a theoretically optimized transonic airfoil
Lores, M. E.; Burdges, K. P.; Shrewsbury, G. D.
1978-01-01
Numerical optimization was used in conjunction with an inviscid, full potential equation, transonic flow analysis computer code to design an upper surface contour for a conventional airfoil to improve its supercritical performance. The modified airfoil was tested in a compressible flow wind tunnel. The modified airfoil's performance was evaluated by comparison with test data for the baseline airfoil and for an airfoil developed by optimization of leading edge of the baseline airfoil. While the leading edge modification performed as expected, the upper surface re-design did not produce all of the expected performance improvements. Theoretical solutions computed using a full potential, transonic airfoil code corrected for viscosity were compared to experimental data for the baseline airfoil and the upper surface modification. These correlations showed that the theory predicted the aerodynamics of the baseline airfoil fairly well, but failed to accurately compute drag characteristics for the upper surface modification.
Transonic Cascade Measurements to Support Analytical Modeling
2007-11-02
RECEIVED JUL 0 12005 FINAL REPORT FOR: AFOSR GRANT F49260-02-1-0284 TRANSONIC CASCADE MEASUREMENTS TO SUPPORT ANALYTICAL MODELING Paul A. Durbin ...PAD); 650-723-1971 (JKE) durbin @vk.stanford.edu; eaton@vk.stanford.edu submitted to: Attn: Dr. John Schmisseur Air Force Office of Scientific Research...both spline and control points for subsequent wall shape definitions. An algebraic grid generator was used to generate the grid for the blade-wall
MACH: Fast Randomized Tensor Decompositions
Tsourakakis, Charalampos E
2009-01-01
Tensors naturally model many real world processes which generate multi-aspect data. Such processes appear in many different research disciplines, e.g, chemometrics, computer vision, psychometrics and neuroimaging analysis. Tensor decompositions such as the Tucker decomposition are used to analyze multi-aspect data and extract latent factors, which capture the multilinear data structure. Such decompositions are powerful mining tools, for extracting patterns from large data volumes. However, most frequently used algorithms for such decompositions involve the computationally expensive Singular Value Decomposition. In this paper we propose MACH, a new sampling algorithm to compute such decompositions. Our method is of significant practical value for tensor streams, such as environmental monitoring systems, IP traffic matrices over time, where large amounts of data are accumulated and the analysis is computationally intensive but also in "post-mortem" data analysis cases where the tensor does not fit in the availa...
50 years of transonic aircraft design
Jameson, Antony; Ou, Kui
2011-07-01
This article traces the evolution of long range jet transport aircraft over the 50 years since Kuechemann founded the journal Progress in Aerospace Sciences. The article is particularly focused on transonic aerodynamics. During Kuechemann's life time a good qualitative understanding had been achieved of transonic flow and swept wing design, but transonic flow remained intractable to quantitative prediction. During the last 50 years this situation has been completely transformed by the introduction of sophisticated numerical algorithms and an astonishing increase in the available computational power, with the consequence that aerodynamic design is now carried out largely by computer simulation. Moreover developments in aerodynamic shape optimization based on control theory enable a competitive swept wing to be designed in just two simulations, as illustrated in the article. While the external appearance of long range jet aircraft has not changed much, advances in information technology have actually transformed the entire design and manufacturing process through parallel advances in computer aided design (CAD), computational structural mechanics (CSM) and multidisciplinary optimization (MDO). They have also transformed aircraft operations through the adoption of digital fly-by-wire and advanced navigational techniques.
Landrum, E. J.; Babb, C. D.
1979-01-01
Flow visualization and force data for a series of six bodies of revolution are presented without analysis. The data were obtained in the Langley Unitary Plan wind tunnel for angles of attack from -4 deg to 60 deg. The Reynolds number used for these tests was 6,600,000 per meter.
Mach stem formation in reflection and focusing of weak shock acoustic pulses.
Karzova, Maria M; Khokhlova, Vera A; Salze, Edouard; Ollivier, Sébastien; Blanc-Benon, Philippe
2015-06-01
The aim of this study is to show the evidence of Mach stem formation for very weak shock waves with acoustic Mach numbers on the order of 10(-3) to 10(-2). Two representative cases are considered: reflection of shock pulses from a rigid surface and focusing of nonlinear acoustic beams. Reflection experiments are performed in air using spark-generated shock pulses. Shock fronts are visualized using a schlieren system. Both regular and irregular types of reflection are observed. Numerical simulations are performed to demonstrate the Mach stem formation in the focal region of periodic and pulsed nonlinear beams in water.
Wornom, Dewey E.
1959-01-01
An investigation of a model of a standard size body in combination with a representative 45 deg swept-wing-fuselage model has been conducted in the Langley 8-foot transonic pressure tunnel over a Mach number range from 0.80 to 1.43. The body, with a fineness ratio of 8.5, was tested with and without fins, and was pylon-mounted beneath the fuselage or wing. Force measurements were obtained on the wing-fuselage model with and without the body, for an angle-of-attack range from -2 deg to approximately 12 deg and an angle-of-sideslip range from -8 deg to 8 deg. In addition, body loads were measured over the same angle-of-attack and angle-of-sideslip range. The Reynolds number for the investigation, based on the wing mean aerodynamic chord, varied from 1.85 x 10(exp 6) to 2.85 x 10(exp 6). The addition of the body beneath the fuselage or the wing increased the drag coefficient of the complete model over the Mach number range tested. On the basis of the drag increase per body, the under-fuselage position was the more favorable. Furthermore, the bodies tended to increase the lateral stability of the complete model. The variation of body loads with angle of attack for the unfinned bodies was generally small and linear over the Mach number range tested with the addition of fins causing large increases in the rates of change of normal-force coefficient and nose-down pitching-moment coefficient. The variation of body side-force coefficient with sideslip for the unfinned body beneath the fuselage was at least twice as large as the variation of this load for the unfinned body beneath the wing. The addition of fins to the body beneath either the fuselage or the wing approximately doubled the rate of change of body side-force coefficient with sideslip. Furthermore, the variation of body side-force coefficient with sideslip for the body beneath the wing was at least twice as large as the variation of this load with angle of attack.
Numerical Investigations of the Benchmark Supercritical Wing in Transonic Flow
Chwalowski, Pawel; Heeg, Jennifer; Biedron, Robert T.
2017-01-01
This paper builds on the computational aeroelastic results published previously and generated in support of the second Aeroelastic Prediction Workshop for the NASA Benchmark Supercritical Wing (BSCW) configuration. The computational results are obtained using FUN3D, an unstructured grid Reynolds-Averaged Navier-Stokes solver developed at the NASA Langley Research Center. The analysis results show the effects of the temporal and spatial resolution, the coupling scheme between the flow and the structural solvers, and the initial excitation conditions on the numerical flutter onset. Depending on the free stream condition and the angle of attack, the above parameters do affect the flutter onset. Two conditions are analyzed: Mach 0.74 with angle of attack 0 and Mach 0.85 with angle of attack 5. The results are presented in the form of the damping values computed from the wing pitch angle response as a function of the dynamic pressure or in the form of dynamic pressure as a function of the Mach number.
Intermittent Flow Regimes in a Transonic Fan Airfoil Cascade
Directory of Open Access Journals (Sweden)
J. Lepicovsky
2004-01-01
velocity.To date, this flow behavior has only been observed in a linear transonic cascade. Further research is necessary to confirm this phenomenon occurs in actual transonic fans and is not the by-product of an endwall restricted linear cascade.
Peterson, Victor L.; Menees, Gene P.
1961-01-01
Tabulated results of a wind-tunnel investigation of the aerodynamic loads on a canard airplane model with twin vertical tails are presented for Mach numbers from 0.70 to 2.22. The Reynolds number for the measurements was 2.9 x 10(exp 6) based on the wing mean aerodynamic chord. The results include local static-pressure coefficients measured on the wing, body, and one of the vertical tails for angles of attack from -4 degrees to 16 degree angles of sideslip of 0 degrees and 5.3 degrees, and nominal canard deflections of O degrees and 10 degrees. Also included are section force and moment coefficients obtained from integrations of the local pressures and model-component force and moment coefficients obtained from integrations of the section coefficients. Geometric details of the model are shown and the locations of the pressure orifices are shown. An index to the data contained herein is presented and definitions of nomenclature are given. Detailed descriptions of the model and experiments and a brief discussion of some of the results are given. Tabulated results of measurements of the aerodynamic loads on the same canard model but having a single vertical tail instead of twin vertical tails are presented.
Peterson, Victor L.; Menees, Gene P.
1961-01-01
Tabulated results of a wind-tunnel investigation of the aerodynamic loads on a canard airplane model with a single vertical tail are presented for Mach numbers from 0.70 to 2.22. The Reynolds number for the measurements was 2.9 x 10(exp 6) based on the wing mean aerodynamic chord. The results include local static pressure coefficients measured on the wing, body, and vertical tail for angles of attack from -4 deg to + 16 deg, angles of sideslip of 0 deg and 5.3 deg, vertical-tail settings of 0 deg and 5 deg, and nominal canard deflections of 0 deg and 10 deg. Also included are section force and moment coefficients obtained from integrations of the local pressures and model-component force and moment coefficients obtained from integrations of the section coefficients. Geometric details of the model and the locations of the pressure orifices are shown. An index to the data contained herein is presented and definitions of nomenclature are given.
Kelly, T. C.
1980-01-01
Pressure and load distributions for a related group of simulated launch vehicle configurations are presented. The configurations were selected so that the nose cone and interstage transition flare components were relatively close to one another and subject to mutual interference effects. Tests extended over a Mach number range from 0.40 to 1.20 at angles of attack from 0 deg to about 10 deg. The test Reynolds numbers, based on main stage diameter, were of the order of 0.00000098.
Muhlstein, L., Jr.; Steinle, F., Jr.
1976-01-01
Fluid dynamic research with the objective of developing new and improved technology in both test facility concepts and test techniques is being reported. A summary of efforts and results thus far obtained in four areas is presented. The four area are: (1) the use of heavy gases to obtain high Reynolds numbers at transonic speeds: (2) high Reynolds number tests of the C-141A wing configuration; (3) performance and flow quality of the pilot injector driven wind tunnel; and (4) integration time required to extract accurate static and dynamic data from tests in transonic wind tunnels. Some of the principal conclusions relative to each of the four areas are: (1) Initial attempts to apply analytical corrections to test results using gases with gamma other than 1.4 to simulate conditions in air show promise but need significant improvement; (2) for the C-141A configuration, no Reynolds number less than the full scale flight value provides an accurate simulation of the full scale flow; (3) high ratios of tunnel mass flow rate to injection mass flow rate and high flow quality can be obtained in an injector driven transonic wind tunnel; and (4) integration times of 0.5 to 1.0 sec may be required for static force and pressure tests, respectively, at some transonic test conditions in order to obtain the required data accuracy.
Interplay between Mach cone and radial expansion in jet events
Tachibana, Y.; Hirano, T.
2016-12-01
We study the hydrodynamic response to jet propagation in the expanding QGP and investigate how the particle spectra after the hydrodynamic evolution of the QGP reflect it. We perform simulations of the space-time evolution of the QGP in gamma-jet events by solving (3+1)-dimensional ideal hydrodynamic equations with source terms. Mach cone is induced by the jet energy deposition and pushes back the radial flow of the expanding background. Especially in the case when the jet passage is off-central one, the number of particles emitted in the direction of the push back decreases. This is the signal including the information about the formation of the Mach cone and the jet passage in the QGP fluid.
Mach, the Universe, and Foundations of Mechanics
Mashhoon, B
2011-01-01
Barbour's response to our recent paper on "Mach's principle and higher-dimensional dynamics" describes an approach to Mach's principle in which the universe as a whole is involved in the definition of inertial frames of reference. Moreover, Barbour's theoretical procedure is in agreement with general relativity for a finite universe that is spatially closed. However, we prefer an operational approach that relies ultimately on observational data.
Emergent gravity of fractons: Mach's principle revisited
Pretko, Michael
2017-07-01
Recent work has established the existence of stable quantum phases of matter described by symmetric tensor gauge fields, which naturally couple to particles of restricted mobility, such as fractons. We focus on a minimal toy model of a rank 2 tensor gauge field, consisting of fractons coupled to an emergent graviton (massless spin-2 excitation). We show how to reconcile the immobility of fractons with the expected gravitational behavior of the model. First, we reformulate the fracton phenomenon in terms of an emergent center of mass quantum number, and we show how an effective attraction arises from the principles of locality and conservation of center of mass. This interaction between fractons is always attractive and can be recast in geometric language, with a geodesiclike formulation, thereby satisfying the expected properties of a gravitational force. This force will generically be short-ranged, but we discuss how the power-law behavior of Newtonian gravity can arise under certain conditions. We then show that, while an isolated fracton is immobile, fractons are endowed with finite inertia by the presence of a large-scale distribution of other fractons, in a concrete manifestation of Mach's principle. Our formalism provides suggestive hints that matter plays a fundamental role, not only in perturbing, but in creating the background space in which it propagates.
Calculation of transonic flow in a linear cascade
Donovan, L. F.
1984-01-01
Turbomachinery blade designs are becoming more aggressive in order to achieve higher loading and greater range. New analysis tools are required to cope with these heavily loaded blades that may operate with a thin separated region near the trailing edge on the suction surface. An existing, viscous airfoil code was adapted to cascade conditions in an attempt to provide this capability. Comparisons with recently obtained data show that calculated and experimental surface Mach numbers were in good agreement but loss coefficients and outlet air angles were not. Previously announced in STAR as N84-24539
Fenomenologia e fenomenismo em Husserl e Mach
Directory of Open Access Journals (Sweden)
Denis Fisette
2009-12-01
Full Text Available Como conciliar as repetidas críticas ao fenomenismo de Mach, um pouco por toda a obra de Husserl, com o papel proeminente que Husserl parece nele reconhecer em seus últimos trabalhos, quanto à gênese de sua própria fenomenologia? Para responder a essa questão, examinaremos, primeiramente, a relação estreita que Husserl estabelece entre o método fenomenológico e o descritivismo de Mach à luz do debate que opõe nativismo e empirismo sobre a origem da percepção do espaço. Em seguida, examinaremos dois aspectos da crítica que Husserl faz ao positivismo de Mach: o primeiro se refere ao fenomenismo e sua doutrina dos elementos, enquanto o segundo, ao princípio de economia de pensamento, que Husserl associa a uma forma de psicologismo em Prolegômenos. A hipótese que nos guiará nesse estudo é que as opiniões aparentemente contraditórias de Husserl sobre o positivismo de Mach se explicam em parte pelo estatuto duplo que a fenomenologia recebe em seus últimos trabalhos: enquanto programa filosófico, ela se opõe explicitamente ao positivismo; enquanto método, ela se aparenta ao descritivismo de Mach. Concluiremos com a ideia de que esses dois filósofos de origem checa perseguiam o objetivo comum de apreender o sentido originário de positividade.How to conciliate the recurrent criticisms to Mach's phenomenism, a bit in all Husserl's work, with the outstanding role Husserl seems to recognise in phenomenism in his last works, as to the genesis of his own phenomenology? In order to answer this question, we examine, first, the close relationship stablished by Husserl between the phenomenological method and Mach's descriptivism in light of the debate that opposes nativism and empiricism regarding the origin of the perception of space. Next, we examine two features of Husserl's criticism to Mach's positivism: the first refers to phenomenism ans its doctrine of elements, and the second, to the principle of economy of thought, which
Numerical Simulation of Shock Bubble Interaction with Different Mach Numbers
Yang, Jie; Wan, Zhen-Hua; Wang, Bo-Fu; Sun, De-Jun
2015-03-01
Not Available Supported by the National Natural Science Foundation of China under Grant Nos 11232011 and 11402262, the 111 Project under Grant No B07033, and the China Postdoctoral Science Foundation Funded Project under Grant No 2014M561833.
Experimental Studies of Very-High Mach Number Hydrodynamics
1994-02-14
intensity of the from Rotman (1991. symbol R) shock amplification of the density fluctuations as a parameter, together with a turbulent kinetic energy...overlapsubgrid scale model while predicting an increase in the model where an algebraic identity provides a procedure for overall grid spectral energy... Rotman , and W. P. improvement in the dissipative near-wal region. Dannevik during the course of this work. Figure 8. indicates that for even steeper
Plasma flow at a high Mach-number
Energy Technology Data Exchange (ETDEWEB)
Yu, Bing; Hameiri, Eliezer [Courant Institute of Mathematical Sciences, New York University New York, New York 10012 (United States)
2013-09-15
Unlike the case of static magnetohydrodynamic (MHD) equilibria, where an expansion in large aspect ratio of toroidal devices is common, cases of MHD equilibria with flow are rarely treated this way, and when this is done the expansion tends to be only partial. The main reason for the difference seems to be the difficulty of expanding the larger system of equilibrium equations with flow. Here, we use a recent expansion technique which employs a variational principle to simplify the process [E. Hameiri, Phys. Plasmas 20, 024504 (2013)]. We treat four cases of MHD equilibria with flow, developing their asymptotic expansions in full, and for an application consider the effect of the flow on the Shafranov shift.
Hydrocarbon-Fueled Scramjet Research at Hypersonic Mach Numbers
2005-03-31
hypersonic flow. Laser-induced fluorescence has the threefold advantages for combustion studies of being non- intrusive , species-specific and highly sensitive...Propulsion Conference and Exhibit, Seattle, WA. Griffiths, A. (2004), Development and Demonstration of a Diode Laser Based Temperature and Water Vapour
Transonic properties of the accretion disk around compact objects
Mukhopadhyay, Banibrata
2008-01-01
An accretion flow is necessarily transonic around a black hole. However, around a neutron star it may or may not be transonic, depending on the inner disk boundary conditions influenced by the neutron star. I will discuss various transonic behavior of the disk fluid in general relativistic (or pseudo general relativistic) framework. I will address that there are four types of sonic/critical point possible to form in an accretion disk. It will be shown that how the fluid properties including location of sonic points vary with angular momentum of the compact object which controls the overall disk dynamics and outflows.
Wu, Chung-Hua
1993-01-01
This report represents a general theory applicable to axial, radial, and mixed flow turbomachines operating at subsonic and supersonic speeds with a finite number of blades of finite thickness. References reflect the evolution of computational methods used, from the inception of the theory in the 50's to the high-speed computer era of the 90's. Two kinds of relative stream surfaces, S(sub 1) and S(sub 2), are introduced for the purpose of obtaining a three-dimensional flow solution through the combination of two-dimensional flow solutions. Nonorthogonal curvilinear coordinates are used for the governing equations. Methods of computing transonic flow along S(sub 1) and S(sub 2) stream surfaces are given for special cases as well as for fully three-dimensional transonic flows. Procedures pertaining to the direct solutions and inverse solutions are presented. Information on shock wave locations and shapes needed for computations are discussed. Experimental data from a Deutsche Forschungs- und Versuchsanstalt fur Luft- und Raumfahrt e.V. (DFVLR) rotor and from a Chinese Academy of Sciences (CAS) transonic compressor rotor are compared with the computed flow properties.
Directory of Open Access Journals (Sweden)
Xiaojing Wu
2017-01-01
Full Text Available Airfoil geometric uncertainty can generate aerodynamic characteristics fluctuations. Uncertainty quantification is applied to compute its impact on the aerodynamic characteristics. In addition, the contribution of each uncertainty variable to aerodynamic characteristics should be computed by the uncertainty sensitivity analysis. In the paper, Sobol’s analysis is used for uncertainty sensitivity analysis and a nonintrusive polynomial chaos method is used for uncertainty quantification and Sobol’s analysis. It is difficult to describe geometric uncertainty because it needs a lot of input parameters. In order to alleviate the contradiction between the variable dimension and computational cost, a principal component analysis is introduced to describe geometric uncertainty of airfoil. Through this technique, the number of input uncertainty variables can be reduced and typical global deformation modes can be obtained. By uncertainty quantification, we can learn that the flow characteristics of shock wave and boundary layer separation are sensitive to the geometric uncertainty in transonic region, which is the main reason that transonic drag is sensitive to the geometric uncertainty. The sensitivity analysis shows that the model can be simplified by eliminating unimportant geometric modes. Moreover, which are the most important geometric modes to transonic aerodynamics can be learnt. This is very helpful for airfoil design.
Force Measurement Improvements to the National Transonic Facility Sidewall Model Support System
Goodliff, Scott L.; Balakrishna, Sundareswara; Butler, David; Cagle, C. Mark; Chan, David; Jones, Gregory S.; Milholen, William E., II
2016-01-01
The National Transonic Facility is a transonic pressurized cryogenic facility. The development of the high Reynolds number semi-span capability has advanced over the years to include transonic active flow control and powered testing using the sidewall model support system. While this system can be used in total temperatures down to -250Â F for conventional unpowered configurations, it is limited to temperatures above -60Â F when used with powered models that require the use of the high-pressure air delivery system. Thermal instabilities and non-repeatable mechanical arrangements revealed several data quality shortfalls by the force and moment measurement system. Recent modifications to the balance cavity recirculation system have improved the temperature stability of the balance and metric model-to-balance hardware. Changes to the mechanical assembly of the high-pressure air delivery system, particularly hardware that interfaces directly with the model and balance, have improved the repeatability of the force and moment measurement system. Drag comparisons with the high-pressure air system removed will also be presented in this paper.
PIV study on a shock-induced separation in a transonic flow
Energy Technology Data Exchange (ETDEWEB)
Sartor, Fulvio; Losfeld, Gilles; Bur, Reynald [ONERA, Meudon (France)
2012-09-15
A transonic interaction between a steady shock wave and a turbulent boundary layer in a Mach 1.4 channel flow is experimentally investigated by means of particle image velocimetry (PIV). In the test section, the lower wall is equipped with a contour profile shaped as a bump allowing flow separation. The transonic interaction, characterized by the existence in the outer flow of a lambda shock pattern, causes the separation of the boundary layer, and a low-speed recirculating bubble is observed downstream of the shock foot. Two-component PIV velocity measurements have been performed using an iterative gradient-based cross-correlation algorithm, providing high-speed and flexible calculations, instead of the classic multi-pass processing with FFT-based cross-correlation. The experiments are performed discussing all the hypotheses linked to the experimental set-up and the technique of investigation such as the two-dimensionality assumption of the flow, the particle response assessment, the seeding system, and the PIV correlation uncertainty. Mean velocity fields are presented for the whole interaction with particular attention for the recirculating bubble downstream of the detachment, especially in the mixing layer zone where the effects of the shear stress are most relevant. Turbulence is discussed in details, the results are compared to previous study, and new results are given for the turbulent production term and the return to isotropy mechanism. Finally, using different camera lens, a zoom in the vicinity of the wall presents mean and turbulent velocity fields for the incoming boundary layer. (orig.)
Chaparro, Daniel; Fujiwara, Gustavo E. C.; Ting, Eric; Nguyen, Nhan
2016-01-01
The need to rapidly scan large design spaces during conceptual design calls for computationally inexpensive tools such as the vortex lattice method (VLM). Although some VLM tools, such as Vorview have been extended to model fully-supersonic flow, VLM solutions are typically limited to inviscid, subcritical flow regimes. Many transport aircraft operate at transonic speeds, which limits the applicability of VLM for such applications. This paper presents a novel approach to correct three-dimensional VLM through coupling of two-dimensional transonic small disturbance (TSD) solutions along the span of an aircraft wing in order to accurately predict transonic aerodynamic loading and wave drag for transport aircraft. The approach is extended to predict flow separation and capture the attenuation of aerodynamic forces due to boundary layer viscosity by coupling the TSD solver with an integral boundary layer (IBL) model. The modeling framework is applied to the NASA General Transport Model (GTM) integrated with a novel control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF).
Transonic Wing Shape Optimization Using a Genetic Algorithm
Holst, Terry L.; Pulliam, Thomas H.; Kwak, Dochan (Technical Monitor)
2002-01-01
A method for aerodynamic shape optimization based on a genetic algorithm approach is demonstrated. The algorithm is coupled with a transonic full potential flow solver and is used to optimize the flow about transonic wings including multi-objective solutions that lead to the generation of pareto fronts. The results indicate that the genetic algorithm is easy to implement, flexible in application and extremely reliable.
Design and Test of a Transonic Axial Splittered Rotor
2015-06-15
AXIAL SPLITTERED ROTOR A new design procedure was developed that uses commercial-off-the-shelf software (MATLAB, SolidWorks , and ANSYS-CFX) for the...TRANSONIC AXIAL SPLITTERED ROTOR Report Title A new design procedure was developed that uses commercial-off-the-shelf software (MATLAB, SolidWorks , and...that uses commercial-off-the-shelf software (MATLAB, SolidWorks , and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial
Demonstration of PIV in a Transonic Compressor
Wernet, Mark P.
1998-01-01
Particle Imaging Velocimetry (PIV) is a powerful measurement technique which can be used as an alternative or complementary approach to Laser Doppler Velocimetry (LDV) in a wide range of research applications. PIV data are measured simultaneously at multiple points in space, which enables the investigation of the non-stationary spatial structures typically encountered in turbomachinery. Many of the same issues encountered in the application of LDV techniques to rotating machinery apply in the application of PIV. Preliminary results from the successful application of the standard 2-D PIV technique to a transonic axial compressor are presented. The lessons learned from the application of the 2-D PIV technique will serve as the basis for applying 3-component PIV techniques to turbomachinery.
Numerical simulation of transonic flows in diffusers
Liou, M.-S.; Coakley, T. J.; Bergmann, M. Y.
1981-01-01
Numerical simulations were made of two-dimensional transonic flows in diffusers, including flow separation induced by a shock or adverse pressure gradient. The mass-averaged, time-dependent, compressible Navier-Stokes equations, simplified by the thin-layer approximation, were solved using MacCormack's hybrid method. The eddy-viscosity formulation was described by the Wilcox-Rubesin's two-equation, k-omega model. Detailed comparison of the computed results with measurements showed good agreement in all cases, including one with massive separation induced by a strong shock. The computation correctly predicted the details of a distinct lambda shock pattern, closely duplicating the configuration observed experimentally in spark-schlieren photographs.
Paryz, Roman W.
2014-01-01
Several upgrade projects have been completed at the NASA Langley Research Center National Transonic Facility over the last 1.5 years in an effort defined as STARBUKS - Subsonic Transonic Applied Refinements By Using Key Strategies. This multi-year effort was undertaken to improve NTF's overall capabilities by addressing Accuracy and Validation, Productivity, and Reliability areas at the NTF. This presentation will give a brief synopsis of each of these efforts.
Energy Technology Data Exchange (ETDEWEB)
Yaga, M.; Nagai, M.; Haga, T. [Univ. of the Ryukyus, Okinawa (Japan). College of Engineering; Miyara, T. [University of the Ryukyus, Okinawa (Japan). Faculty of Medicine; Tomita, n. [Hitachi, Ltd., Tokyo (Japan)
1996-03-25
The effects of a passive control boundary layer applied around the throat on a transonic diffuser were investigated experimentally by wall static pressure measurements and by schlieren optical observations. The experiments were conducted using three kinds of the diffuser walls, one was a solid wall and the other two were porous with a 30 mm or 50 mm-long cavity underneath enabling the flow around the shock wave to circulate through the porous wall. The results show that the Mach waves normal to the flow were observed when diffusers were almost choked and that the pressure fluctuations in a transonic diffuser were greatly reduced by passive control. According to the frequency analysis, the frequency range attenuated by passive control is between about 700 Hz and 1 kHz regardless of the length of the cavities. 15 refs., 7 figs.
Kulkarni, Sameer; Beach, Timothy A.; Jorgenson, Philip C.; Veres, Joseph P.
2017-01-01
A 24 foot diameter 3-stage axial compressor powered by variable-speed induction motors provides the airflow in the closed-return 11- by 11-Foot Transonic Wind Tunnel (11-Foot TWT) Facility at NASA Ames Research Center at Moffett Field, California. The facility is part of the Unitary Plan Wind Tunnel, which was completed in 1955. Since then, upgrades made to the 11-Foot TWT such as flow conditioning devices and instrumentation have increased blockage and pressure loss in the tunnel, somewhat reducing the peak Mach number capability of the test section. Due to erosion effects on the existing aluminum alloy rotor blades, fabrication of new steel rotor blades is planned. This presents an opportunity to increase the Mach number capability of the tunnel by redesigning the compressor for increased pressure ratio. Challenging design constraints exist for any proposed design, demanding the use of the existing driveline, rotor disks, stator vanes, and hub and casing flow paths, so as to minimize cost and installation time. The current effort was undertaken to characterize the performance of the existing compressor design using available design tools and computational fluid dynamics (CFD) codes and subsequently recommend a new compressor design to achieve higher pressure ratio, which directly correlates with increased test section Mach number. The constant cross-sectional area of the compressor leads to highly diffusion factors, which presents a challenge in simulating the existing design. The CFD code APNASA was used to simulate the aerodynamic performance of the existing compressor. The simulations were compared to performance predictions from the HT0300 turbomachinery design and analysis code, and to compressor performance data taken during a 1997 facility test. It was found that the CFD simulations were sensitive to endwall leakages associated with stator buttons, and to a lesser degree, under-stator-platform flow recirculation at the hub. When stator button leakages were
Gross, A. R.; Steinle, F. W., Jr.
1975-01-01
A NACA 64A010 pressure-instrumented airfoil was tested at transonic speeds over a range of angle of attack from -1 to 12 degrees at various Reynolds numbers ranging from 2 to 6 million in air, argon, Freon 12, and a mixture of argon and Freon 12 having a ratio of specific heats corresponding to air. Good agreement of results is obtained for conditions where compressibility is not significant and for the air and comparable argon-Freon 12 mixture. Comparison of heavy gas results with air, when adjusted for transonic similarity, show improved, but less than desired agreement.
Mach bands change asymmetrically during solar eclipses.
Ross, John; Diamond, Mark R; Badcock, David R
2003-01-01
Observations made during two partial eclipses of the Sun show that the Mach bands on shadows cast by the Sun disappear and reappear asymmetrically as an eclipse progresses. These changes can be explained as due to changes in the shape of the penumbras of shadows as the visible portion of the Sun forms crescents of different orientation.
Edge and divertor plasma measurements with ion sensitive and Mach probes in LHD
Energy Technology Data Exchange (ETDEWEB)
Hayashi, Y., E-mail: shihaya_uki884@yahoo.co.jp [Nagano National College of Technology, 716 Tokuma, Nagano 381-8550 (Japan); Ezumi, N. [Nagano National College of Technology, 716 Tokuma, Nagano 381-8550 (Japan); Masuzaki, S.; Tanaka, H.; Kobayashi, M. [National Institute for Fusion Science, 322-6 Oroshi-cho, Toki 509-5292 (Japan); Sawada, K. [Shinshu University, Wakasato, Nagano 380-8553 (Japan); Ohno, N. [Nagoya University, Furo-cho Chikusa-ku, Nagoya 464-8603 (Japan)
2013-07-15
Spatial profiles of plasma flow and Mach number in the stochastic magnetic boundary layer as well as ion temperature (T{sub i}) and electron temperature (T{sub e}) in the divertor region in Large Helical Device (LHD) have been studied by a movable multiple functions probe, which consists of Mach probes and an ion sensitive probe. The results of ion saturation current measurements indicated plasma flow direction is alternated in the stochastic magnetic boundary. Mach number profiles for different plasma densities have been evaluated experimentally which compared with 3-D transport code. T{sub i} and T{sub e} in the divertor region measured by the ion sensitive probe decreased with increasing line-averaged density. Although T{sub i} was higher than T{sub e} in the low density plasma, both temperatures became almost the same at higher density.
Periodic transonic flow simulation using fourier-based algorithm
Energy Technology Data Exchange (ETDEWEB)
Mohaghegh, Mohammad Reza [Islamic Azad University, Tehran (Iran, Islamic Republic of); Malekjafarian, Majid [University of Birjand, Birjand (Iran, Islamic Republic of)
2014-10-15
The present research simulates time-periodic unsteady transonic flow around pitching airfoils via the solution of unsteady Euler and Navier-Stokes equations, using time spectral method (TSM) and compares it with the traditional methods like BDF and explicit structured adaptive grid method. The TSM uses a Fourier representation in time and hence solves for the periodic state directly without resolving transients (which consume most of the resources in a time-accurate scheme). Mathematical tools used here are discrete Fourier transformations. The TSM has been validated with 2D external aerodynamics test cases. These test cases are NACA 64A010 (CT6) and NACA 0012 (CT1 and CT5) pitching airfoils. Because of turbulent nature of flow, Baldwin-Lomax turbulence model has been used in viscous flow analysis with large oscillation amplitude (CT5 type). The results presented by the TSM are compared with experimental data and the two other methods. By enforcing periodicity and using Fourier representation in time that has a spectral accuracy, tremendous reduction of computational cost has been obtained compared to the conventional time-accurate methods. Results verify the small number of time intervals per pitching cycle (just four time intervals) required to capture the flow physics with small oscillation amplitude (CT6) and large oscillation amplitude (CT5) as compared to the other two methods.
Inlet Turbulence and Length Scale Measurements in a Large Scale Transonic Turbine Cascade
Thurman, Douglas; Flegel, Ashlie; Giel, Paul
2014-01-01
Constant temperature hotwire anemometry data were acquired to determine the inlet turbulence conditions of a transonic turbine blade linear cascade. Flow conditions and angles were investigated that corresponded to the take-off and cruise conditions of the Variable Speed Power Turbine (VSPT) project and to an Energy Efficient Engine (EEE) scaled rotor blade tip section. Mean and turbulent flowfield measurements including intensity, length scale, turbulence decay, and power spectra were determined for high and low turbulence intensity flows at various Reynolds numbers and spanwise locations. The experimental data will be useful for establishing the inlet boundary conditions needed to validate turbulence models in CFD codes.
Mach-Zehnder interferometer for movement monitoring
Vasinek, Vladimir; Cubik, Jakub; Kepak, Stanislav; Doricak, Jan; Latal, Jan; Koudelka, Petr
2012-06-01
Fiber optical interferometers belong to highly sensitive equipments that are able to measure slight changes like distortion of shape, temperature and electric field variation and etc. Their great advantage is that they are insensitive on ageing component, from which they are composed of. It is in virtue of herewith, that there are evaluated no changes in optical signal intensity but number interference fringes. To monitor the movement of persons, eventually to analyze the changes in state of motion we developed method based on analysis the dynamic changes in interferometric pattern. We have used Mach- Zehnder interferometer with conventional SM fibers excited with the DFB laser at wavelength of 1550 nm. It was terminated with optical receiver containing InGaAs PIN photodiode. Its output was brought into measuring card module that performs on FFT of the received interferometer signal. The signal rises with the composition of two waves passing through single interferometer arm. The optical fiber SMF 28e in one arm is referential; the second one is positioned on measuring slab at dimensions of 1x2m. A movement of persons around the slab was monitored, signal processed with FFT and frequency spectra were evaluated. They rose owing to dynamic changes of interferometric pattern. The results reflect that the individual subjects passing through slab embody characteristic frequency spectra, which are individual for particular persons. The scope of measuring frequencies proceeded from zero to 10 kHz. It was also displayed in experiments that the experimental subjects, who walked around the slab and at the same time they have had changed their state of motion (knee joint fixation), embodied characteristic changes in their frequency spectra. At experiments the stability of interferometric patterns was evaluated as from time aspects, so from the view of repeated identical experiments. Two kinds of balls (tennis and ping-pong) were used to plot the repeatability measurements and
The Effect of Ultrapolish on a Transonic Axial Rotor
Roberts, William B.; Thorp, Scott; Prahst, Patricia S.; Strazisar, Anthony
2005-01-01
Back-to-back testing has been done using NASA fan rotor 67 in the Glenn Research Center W8 Axial Compressor Test Facility. The rotor was baseline tested with a normal industrial RMS surface finish of 0.5-0.6 m (20-24 microinches) at 60, 80 and 100% of design speed. At design speed the tip relative Mach number was 1.38. The blades were then removed from the facility and ultrapolished to a surface finish of 0.125 m (5 microinch) or less and retested. At 100% speed near the design point, the ultrapolished blades showed approximately 0.3 - 0.5% increase in adiabatic efficiency. The difference was greater near maximum flow. Due to increased relative measurement error at 60 and 80% speed, the performance difference between the normal and ultrapolished blades was indeterminate at these speeds.
Applications of a transonic wing design method
Campbell, Richard L.; Smith, Leigh A.
1989-01-01
A method for designing wings and airfoils at transonic speeds using a predictor/corrector approach was developed. The procedure iterates between an aerodynamic code, which predicts the flow about a given geometry, and the design module, which compares the calculated and target pressure distributions and modifies the geometry using an algorithm that relates differences in pressure to a change in surface curvature. The modular nature of the design method makes it relatively simple to couple it to any analysis method. The iterative approach allows the design process and aerodynamic analysis to converge in parallel, significantly reducing the time required to reach a final design. Viscous and static aeroelastic effects can also be accounted for during the design or as a post-design correction. Results from several pilot design codes indicated that the method accurately reproduced pressure distributions as well as the coordinates of a given airfoil or wing by modifying an initial contour. The codes were applied to supercritical as well as conventional airfoils, forward- and aft-swept transport wings, and moderate-to-highly swept fighter wings. The design method was found to be robust and efficient, even for cases having fairly strong shocks.
Evolution of transonicity in an accretion disc
Ray, A K; Ray, Arnab K.; Bhattacharjee, Jayanta K.
2007-01-01
For inviscid, rotational accretion flows driven by a general pseudo-Newtonian potential on to a Schwarzschild black hole, the only possible fixed points are saddle points and centre-type points. For the specific choice of the Newtonian potential, the flow has only two critical points, of which the outer one is a saddle point while the inner one is a centre-type point. A restrictive upper bound is imposed on the admissible range of values of the angular momentum of sub-Keplerian flows through a saddle point. These flows are very unstable to any deviation from a necessarily precise boundary condition. The difficulties against the physical realisability of a solution passing through the saddle point have been addressed through a temporal evolution of the flow, which gives a non-perturbative mechanism for selecting a transonic solution passing through the saddle point. An equation of motion for a real-time perturbation about the stationary flows reveals a very close correspondence with the metric of an acoustic b...
On Mach's critique of Newton and Copernicus
Hartman, H I; Hartman, Herbert I.; Nissim-Sabat, Charles
2003-01-01
Maintaining the relativity of all motion, especially rotational motion, Mach denied the existence of absolute motion and absolute space. He maintained the equivalence of the Ptolemaic and the Copernican systems and the equivalence of a fixed bucket in a rotating universe with the converse. An analysis of the Foucault pendulum shows that there cannot be a fixed bucket in a rotating universe. Also, Mach's views violate the physics he espoused: non-inertial experiments, e.g. stellar aberration and electromagnetic effects, distinguish between a rotating bucket in a fixed universe and the converse, between the Copernican and the Ptolemaic systems, and establish that one cannot ascribe all observations solely to relative motion between a system and the universe.
Finite elements and finite differences for transonic flow calculations
Hafez, M. M.; Murman, E. M.; Wellford, L. C.
1978-01-01
The paper reviews the chief finite difference and finite element techniques used for numerical solution of nonlinear mixed elliptic-hyperbolic equations governing transonic flow. The forms of the governing equations for unsteady two-dimensional transonic flow considered are the Euler equation, the full potential equation in both conservative and nonconservative form, the transonic small-disturbance equation in both conservative and nonconservative form, and the hodograph equations for the small-disturbance case and the full-potential case. Finite difference methods considered include time-dependent methods, relaxation methods, semidirect methods, and hybrid methods. Finite element methods include finite element Lax-Wendroff schemes, implicit Galerkin method, mixed variational principles, dual iterative procedures, optimal control methods and least squares.
Static and dynamic aspects of transonicity in Bondi accretion
Ray, A K; Ray, Arnab K.; Bhattacharjee, Jayanta K.
2006-01-01
Transonicity in a spherically symmetric accreting system has been considered in both the stationary and the dynamic regimes. The stationary flow, set up as a dynamical system, has been shown to be greatly unstable to even the minutest possible deviation in the boundary condition for transonicity. With the help of a simple analytical model, and some numerical modelling, it has then been argued that the flow indeed becomes transonic and stable, when the evolution of the flow is followed through time. The time-dependent approach also shows that there is a remarkable closeness between an equation of motion for a perturbation in the flow, and the metric of an analog acoustic black hole.
Dynamics of compressional Mach cones in a strongly coupled complex plasma
Bandyopadhyay, P; Kadyan, Sangeeta; Sen, Abhijit
2016-01-01
Using a Generalised-Hydrodynamic (GH) fluid model we study the influence of strong coupling induced modification of the fluid compressibility on the dynamics of compressional Mach cones in a dusty plasma medium. A significant structural change of lateral wakes for a given Mach number and Epstein drag force is found in the strongly coupled regime. With the increase of fluid compressibility, the peak amplitude of the normalised perturbed dust density first increases and then decreases monotonically after reaching its maximum value. It is also noticed that the opening angle of the cone structure decreases with the increase of the compressibility of the medium and the arm of the Mach cone breaks up into small structures in the velocity vector profile when the coupling between the dust particles increases.
Institute of Scientific and Technical Information of China (English)
履之
1994-01-01
Most engines compress air, add fuel and burn it, and then allow theheated gas to expand, creating power or thrust. A radical aircraft enginedevised by ONERA, France’s equivalent of NASA, does the opposite.The Priam inverse-cycle" engine is designed for hypersonic speedsabove Mach 4 (2, 650 mph). Conventional jets do not work at suchspeeds, because the air becomes so hot when it is rammed into the
Dynamic transition from Mach to regular reflection of shock waves in a steady flow
CSIR Research Space (South Africa)
Naidoo, K
2014-07-01
Full Text Available decreased with increased rotation speed. The sensitivity of the transition angle to changing the rotation point from the trailing edge to the experimental model pivot point was investigated briefly at a free-stream Mach number of M=2.98 with M(subE)=-0...
Mach's Principle and Higher-Dimensional Dynamics
Mashhoon, B
2011-01-01
We briefly discuss the current status of Mach's principle in general relativity and point out that its last vestige, namely, the gravitomagnetic field associated with rotation, has recently been measured for the earth in the GP-B experiment. Furthermore, in his analysis of the foundations of Newtonian mechanics, Mach provided an operational definition for inertial mass and pointed out that time and space are conceptually distinct from their operational definitions by means of masses. Mach recognized that this circumstance is due to the lack of any a priori connection between the inertial mass of a body and its Newtonian state in space and time. One possible way to improve upon this situation in classical physics is to associate mass with an extra dimension. Indeed, Einstein's theory of gravitation can be locally embedded in a Ricci-flat 5D manifold such that the 4D energy-momentum tensor appears to originate from the existence of the extra dimension. An outline of such a 5D Machian extension of Einstein's gen...
Institute of Scientific and Technical Information of China (English)
向宏辉; 侯敏杰; 梁俊; 葛宁; 刘志刚
2015-01-01
The reasonable suction position was determined according to conventional transonic diffuser cas⁃cade experimental results. The boundary layer suction experiments of the cascade were carried out under different operating conditions. The changes of blade surface Mach number distribution, outlet wake and to⁃tal pressure loss coefficient which caused by boundary layer suction effect were analyzed. The results indi⁃cated that the impact of slotting on the original flow field structure of conventional cascade was small. But when the suction slot was located in the peak Mach number position, it would produce effects on the down⁃stream flow of the cascade. The suction in proper position can suppress surface flow separation of the tran⁃sonic cascade, and only the suction flow rate reached a certain value, the boundary layer suction had signifi⁃cant positive effects on the cascade aerodynamic performance. The total pressure loss coefficient of the tran⁃sonic cascade decreased 7.8%when the suction flow rate was 0.87%.%基于常规跨声速扩压叶栅吹风试验结果确定合理抽吸位置，并在此基础上对该叶栅进行多种工况的附面层抽吸试验，分析附面层抽吸作用下叶片表面马赫数、出口尾迹与总压损失系数的变化。结果表明：开设抽吸缝对常规跨声速叶栅原有流场结构的总体影响较小，但当抽吸缝位于马赫数峰值位置时，会对下游流动产生一定扰动。在适当位置抽吸能抑制跨声速叶栅表面流动分离，且只有抽吸量达到一定数值后，附面层抽吸作用才会对叶栅气动性能起到明显正效果。当抽吸量达到0.87%时，该跨声速叶栅总压损失系数降低了7.8%。
Institute of Scientific and Technical Information of China (English)
T. Irie; T. Yasunobu; H. Kashimura; T. Setoguchi
2003-01-01
When the high-pressure gas is exhausted to the vacuum chamber from the nozzle, the underexpanded supersonic jet contained with the Mach disk is generally formed. The eventual purpose of this study is to clarify the unsteady phenomenon of the underexpanded free jet when the back pressure continuously changes with time. The characteristic of the Mach disk has been clarified in consideration of the diameter and position of it by the numerical analysis in this paper. The sonic jet of the exit Mach number Me=1 is assumed and the axisymmetric conservational equation is solved by the TVD method in the numerical calculation.The diameter and position of the Mach disk differs with the results of a steady jet and the influence on the continuously changing of the back pressure is evidenced from the comparison with the case of steady supersonic jet.
Combustion-Powered Actuation for Dynamic Stall Suppression - Simulations and Low-Mach Experiments
Matalanis, Claude G.; Min, Byung-Young; Bowles, Patrick O.; Jee, Solkeun; Wake, Brian E.; Crittenden, Tom; Woo, George; Glezer, Ari
2014-01-01
An investigation on dynamic-stall suppression capabilities of combustion-powered actuation (COMPACT) applied to a tabbed VR-12 airfoil is presented. In the first section, results from computational fluid dynamics (CFD) simulations carried out at Mach numbers from 0.3 to 0.5 are presented. Several geometric parameters are varied including the slot chordwise location and angle. Actuation pulse amplitude, frequency, and timing are also varied. The simulations suggest that cycle-averaged lift increases of approximately 4% and 8% with respect to the baseline airfoil are possible at Mach numbers of 0.4 and 0.3 for deep and near-deep dynamic-stall conditions. In the second section, static-stall results from low-speed wind-tunnel experiments are presented. Low-speed experiments and high-speed CFD suggest that slots oriented tangential to the airfoil surface produce stronger benefits than slots oriented normal to the chordline. Low-speed experiments confirm that chordwise slot locations suitable for Mach 0.3-0.4 stall suppression (based on CFD) will also be effective at lower Mach numbers.
Numerical simulation of Mach reflection of cellular detonations
Li, J.; Lee, J. H. S.
2016-09-01
The Mach reflection of cellular detonation waves on a wedge is investigated numerically in an attempt to elucidate the effect of cellular instabilities on Mach reflection, the dependence of self-similarity on the thickness of a detonation wave, and the initial development of the Mach stem near the wedge apex. A two-step chain-branching reaction model is used to give a thermally neutral induction zone followed by a chemical reaction zone for the detonation wave. A sufficiently large distance of travel of the Mach stem is computed to observe the asymptotic behavior in the far field. Depending on the scale at which the Mach reflection process occurs, it is found that the Mach reflection of a cellular detonation behaves essentially in the same way as a planar ZND detonation wave. The cellular instabilities, however, cause the triple-point trajectory to fluctuate. The fluctuations are due to interactions of the triple point of the Mach stem with the transverse waves of cellular instabilities. In the vicinity of the wedge apex, the Mach reflection is found to be self-similar and corresponds to that of a shock wave of the same strength, since the Mach stem is highly overdriven initially. In the far field, the triple-point trajectory approaches a straight line, indicating that the Mach reflection becomes self-similar asymptotically. The distance of the approach to self-similarity is found to decrease rapidly with decreasing thickness of the detonation front.
Analysis of some interference effects in a transonic wind tunnel
CSIR Research Space (South Africa)
Lombardi, G
1995-05-01
Full Text Available The effects of the walls of a test section on a model in transonic flow were investigated by using the AGARD Calibration Model B. Tests were carried out in a closed-circuit pressurized tunnel, with a confined square test section of 1.5 m width...
Transonic Pressure-- Sensing Studies Using Drop Test Vehicles
Energy Technology Data Exchange (ETDEWEB)
Pepper, W.B., Jr. [Organization 5141
1954-05-01
Free-flight drop vehicle tests have been made to investigate devices for measuring ambient pressure in the vicinity of a high-fineness-ratio weapon shape throughout the transonic speed range. Various types of nose probes and trailing probes were tested.
Pressure drop and thrust predictions for transonic micronozzle flows
Gomez, J.; Groll, R.
2016-02-01
In this paper, the expansion of xenon, argon, krypton, and neon gases through a Laval nozzle is studied experimentally and numerically. The pressurized gases are accelerated through the nozzle into a vacuum chamber in an attempt to simulate the operating conditions of a cold-gas thruster for attitude control of a micro-satellite. The gases are evaluated at several mass flow rates ranging between 0.178 mg/s and 3.568 mg/s. The Re numbers are low (8-256) and the estimated values of Kn number lie between 0.33 and 0.02 (transition and slip-flow regime). Direct Simulation Monte Carlo (DSMC) and continuum-based simulations with a no-slip boundary condition are performed. The DSMC and the experimental results show good agreement in the range Kn > 0.1, while the Navier-Stokes results describe the experimental data more accurately for Kn gas-independent accommodation coefficients. The thrust delivered by the cold-gas thruster and the specific impulse is determined based on the numerical results. Furthermore, an increase of the thickness of the viscous boundary layer through the diffuser of the micronozzle is observed. This results in a shock-less decrease of the Mach number and the flow velocity, which penalizes thrust efficiency. The negative effect of the viscous boundary layer on thrust efficiency can be lowered through higher values of Re and a reduction of the diffuser length.
Directory of Open Access Journals (Sweden)
Rio Melvin Aro.T
2015-05-01
Full Text Available Flutter is an unstable oscillation which can lead to destruction. Flutter can occur on fixed surfaces, such as blades, wing or the stabilizer. By self-excited aeroelastic instability, flutter can lead to mechanical or structural failure of aircraft engine blades. The modern engines have been designed with increased pressure ratio and reduced weight in order to improve aerodynamic efficiency, resulting in severe aeroelastic problems. Particularly flutter in axial compressors with transonic flow can be characterized by a number of aerodynamic nonlinear effects such as shock boundary layer interaction, rotating stall, and tip vortex instability. Rotating blades operating under high centrifugal forces may also encounter structural nonlinearities due to friction damping and large deformations. In the future work a standard axial flow compressor blade will be taken for analysis, both Subsonic and Transonic range are taken for analysis. Fluid and Structure are two different domains which will be coupled by full system coupling technique to predict the fluttering effect on the compressor blade. ANSYS is a commercial simulation tool, which will be deployed in this work to perform FSI (Fluid Structure Interaction and FSI coupled Modal to predict the flutter in the compressor blades
Computational Design and Analysis of a Transonic Natural Laminar Flow Wing for a Wind Tunnel Model
Lynde, Michelle N.; Campbell, Richard L.
2017-01-01
A natural laminar flow (NLF) wind tunnel model has been designed and analyzed for a wind tunnel test in the National Transonic Facility (NTF) at the NASA Langley Research Center. The NLF design method is built into the CDISC design module and uses a Navier-Stokes flow solver, a boundary layer profile solver, and stability analysis and transition prediction software. The NLF design method alters the pressure distribution to support laminar flow on the upper surface of wings with high sweep and flight Reynolds numbers. The method addresses transition due to attachment line contamination/transition, Gortler vortices, and crossflow and Tollmien-Schlichting modal instabilities. The design method is applied to the wing of the Common Research Model (CRM) at transonic flight conditions. Computational analysis predicts significant extents of laminar flow on the wing upper surface, which results in drag savings. A 5.2 percent scale semispan model of the CRM NLF wing will be built and tested in the NTF. This test will aim to validate the NLF design method, as well as characterize the laminar flow testing capabilities in the wind tunnel facility.
Institute of Scientific and Technical Information of China (English)
熊健; 李国帅; 周强; 李平; 马护生; 王红彪; 刘祥; 黄辉
2016-01-01
In recent decades,PSP (Pressure Sensitive Paint )measurement technique has been improved continuously and applied to engineering wind tunnel testing of two meter scale in many aerodynamic test institutions in the world gradually,for model surface pressure measure-ment,flow visualization on model surface and CFD result validation.A binary PSP measurement system with multi light sources and CCD cameras has been established in the 2.4m transonic wind tunnel,and the many problems affecting measurement accuracy and reliability in large intermittent transonic wind tunnel tests have been solved,such as temperature changing on mod-el surface,illumination uniformity and intensity changing,model vibration,as well as test data correction,spraying and coating calibration,and so on.Meanwhile it has been employed to the surface pressure distribution measurement tests of the large aircraft model and the NASA delta wing model.Test results indicate that PSP coating has little influence on model surface pressure distribution.The difference of the root mean square of Cp between PSP and conventional pressure taps is below 0.03 under the test conditions with Mach numbers ranging from 0.4 to 1.2 and angles of attack varying from -4°to 4°,which is equal to the measurements accuracy of continu-ous transonic wind tunnel of the same scale abroad.In a word,the new PSP system and tech-nique application can provide a new advanced measurement method for the configuration optimum design of flight vehicles and aerodynamics research.%近十几年来,由于压敏漆(Pressure Sensitive Paint、PSP)测量技术的不断完善与发展,国际上主要空气动力试验机构逐步将其应用于2 m量级工程型风洞,完成模型表面压力测量、模型表面流动显示与 CFD 结果验证。在2.4m跨声速风洞建立了双组份、多光源和多CCD的PSP测量系统,解决了大型暂冲式跨超声速风洞试验存在的模型表面温度变化、光照均匀性与强度变化,以及模
On Mach's principle: Inertia as gravitation
Martín, J; Tiemblo, A; Ranada, Antonio F.
2007-01-01
In order to test the validity of Mach's principle, we calculate the action of the entire universe on a test mass in its rest frame, which is an acceleration ${\\bf g}^*$. We show the dependence of the inertia principle on the lapse and the shift. Using the formalism of linearized gravitation, we obtain the non-relativistic limit of ${\\bf g}^*$ in terms of two integrals. We follow then two approaches. In the first one, these integrals are calculated in the actual time section $t=t_0$ up to the distance $R_U=ct_0$. In the more exact and satisfactory second approach, they are calculated over the past light cone using the formalism of the retarded potentials. The aim is to find whether the acceleration $\\dot{\\bf v}$ in the LHS of Newton's second law can be interpreted as a reactive acceleration, in other words, as minus the acceleration of gravity ${\\bf g}^*$ in the rest frame of the accelerated particle ({\\it i. e.} to know whether or not ${\\bf g}^*=-\\dot{\\bf v}$). The results strongly support Mach's idea since t...
Experimental investigation of boundary layer suction effects on a transonic cascade%跨声速叶栅叶表附面层抽吸效应试验
Institute of Scientific and Technical Information of China (English)
王掩刚; 国睿; 任思源; 姜正礼
2011-01-01
以某跨声速、吸附式叶栅为研究对象,在暂冲式叶栅风洞上对其进行了多个状态的吹风试验,对比分析了在通道存在激波条件下,激波前、后抽吸对叶栅性能以及附面层的影响效应.研究结果表明：激波前抽吸使得抽吸缝局部马赫数增大,恶化叶栅性能;激波前抽吸对于来流高亚声和超声速的叶栅损失系数影响趋势一致,随着抽吸系数增加损失系数增加,并且当抽吸系数大于0.2%时,损失系数增加较快;波后抽吸可明显改善叶栅性能,抽吸量越大,抽吸正效应越明显,相比于未抽吸条件,抽吸系数为0.8%时损失系数降低8%、总压恢复系数提高5%.%To explore the effects of boundary layer suction（BLS） on a transonic cascade,several experiments were carried out on an intermittent cascade wind tunnel.The results of different locations in relation to the point for passage shock wave projection of various suction coefficients have been investigated.Experimental results show that：when the location of BLS slot is in front of the point for passage shock wave projection,the local Mach number around the suction slot is increased and cascade performance is deteriorated.The loss coefficient increases with the enhancement of suction coefficient for both higher subsonic and supersonic inlet flow conditions,and when the suction coefficient is greater than 0.2%,the loss coefficient increases rapidly.When BLS slot is located behind the shock wave,the cascade performance is significantly improved,and the positive effect of BLS increases by the suction coefficient.Compared with the condition without suction,the loss coefficient reduces by 8% and the total pressure recovery coefficient increases by 5% when suction coefficient is 0.8%.
Status of advanced airfoil tests in the Langley 0.3-meter transonic cryogenic tunnel
Ladson, C. L.; Ray, E. J.
1984-01-01
A joint NASA/U.S. industry program to test advanced technology airfoils in the Langley 0.3-meter Transonic Tunnel (TCT) was formulated under the Langley ACEE Project Office. The objectives include providing U.S. industry an opportunity to compare their most advanced airfoils to the latest NASA designs by means of high Reynolds number tests in the same facility. At the same time, industry would again experience in the design and construction of cryogenic test techniques. The status and details of the test program are presented. Typical aerodynamic results obtained, to date, are presented at chord Reynolds number up to 45 x 10(6) and are compared to results from other facilities and theory. Details of a joint agreement between NASA and the Deutsche Forschungs- und Versuchsantalt fur Luft- and Raumfahrt e.V. (DFVLR) for tests of two airfoils are also included. Results of these tests will be made available as soon as practical.
Serpentine Diffuser Performance with Emphasis on Future Introduction to a Transonic Fan (Postprint)
2013-01-01
AFRL-RQ-WP-TP-2014-0168 SERPENTINE DIFFUSER PERFORMANCE WITH EMPHASIS ON FUTURE INTRODUCTION TO A TRANSONIC FAN (POSTPRINT) Chase A...June 2012 – 01 December 2012 4. TITLE AND SUBTITLE SERPENTINE DIFFUSER PERFORMANCE WITH EMPHASIS ON FUTURE INTRODUCTION TO A TRANSONIC FAN...resulting in an overall recovery factor of 0.983. Distortion descriptors are presented and discussed. 15. SUBJECT TERMS diffuser , transonic fan
Mach 5 to 7 RBCC Propulsion System Testing at NASA-LeRC HTF
Perkins, H. Douglas; Thomas, Scott R.; Pack, William D.
1996-01-01
A series of Mach 5 to 7 freejet tests of a Rocket Based Combined Cycle (RBCC) engine were cnducted at the NASA Lewis Research Center (LERC) Hypersonic Tunnel Facility (HTF). This paper describes the configuration and operation of the HTF and the RBCC engine during these tests. A number of facility support systems are described which were added or modified to enhance the HTF test capability for conducting this experiment. The unfueled aerodynamic perfor- mance of the RBCC engine flowpath is also presented and compared to sub-scale test results previously obtained in the NASA LERC I x I Supersonic Wind Tunnel (SWT) and to Computational Fluid Dynamic (CFD) analysis results. This test program demonstrated a successful configuration of the HTF for facility starting and operation with a generic RBCC type engine and an increased range of facility operating conditions. The ability of sub-scale testing and CFD analysis to predict flowpath performance was also shown. The HTF is a freejet, blowdown propulsion test facility that can simulate up to Mach 7 flight conditions with true air composition. Mach 5, 6, and 7 facility nozzles are available, each with an exit diameter of 42 in. This combination of clean air, large scale, and Mach 7 capabilities is unique to the HTF. This RBCC engine study is the first engine test program conducted at the HTF since 1974.
Institute of Scientific and Technical Information of China (English)
无
2010-01-01
The pressure ratio required for a turbocharger centrifugal compressor increases with internal combustion engine power density. High pressure ratio causes a transonic flow field at the impeller inducer. Transonic flow narrows the stable flow range and de-teriorates stage efficiency. In this work, an advanced high pressure ratio transonic compressor was designed. The experimental results show that the maximum pressure ratio of this turbocharger is about 4.2, the maximum efficiency is above 80% and the stable flow range at the designed rotating speed is up to 34%. A turbocharger with this transonic compressor has been applied to some vehicle research actually, and improved power density by 40%.
Computational aspects of the prediction of multidimensional transonic flows in turbomachinery
Oliver, D. A.; Sparis, P.
1975-01-01
The analytical prediction and description of transonic flow in turbomachinery is complicated by three fundamental effects: (1) the fluid equations describing the transonic regime are inherently nonlinear, (2) shock waves may be present in the flow, and (3) turbomachine blading is geometrically complex, possessing large amounts of curvature, stagger, and twist. A three-dimensional computation procedure for the study of transonic turbomachine fluid mechanics is described. The fluid differential equations and corresponding difference operators are presented, the boundary conditions for complex blade shapes are described, and the computational implementation and mapping procedures are developed. Illustrative results of a typical unthrottled transonic rotor are also presented.
Ernst Mach and the episode of the monocular depth sensations.
Banks, E C
2001-01-01
Although Ernst Mach is widely recognized in psychology for his discovery of the effects of lateral inhibition in the retina ("Mach Bands"), his contributions to the theory of depth perception are not as well known. Mach proposed that steady luminance gradients triggered sensations of depth. He also expanded on Ewald Hering's hypothesis of "monocular depth sensations," arguing that they were subject to the same principle of lateral inhibition as light sensations were. Even after Hermann von Helmholtz's attack on Hering in 1866, Mach continued to develop theories involving the monocular depth sensations, proposing an explanation of perspective drawings in which the mutually inhibiting depth sensations scaled to a mean depth. Mach also contemplated a theory of stereopsis in which monocular depth perception played the primary role. Copyright 2001 John Wiley & Sons, Inc.
A finite-difference method for transonic airfoil design.
Steger, J. L.; Klineberg, J. M.
1972-01-01
This paper describes an inverse method for designing transonic airfoil sections or for modifying existing profiles. Mixed finite-difference procedures are applied to the equations of transonic small disturbance theory to determine the airfoil shape corresponding to a given surface pressure distribution. The equations are solved for the velocity components in the physical domain and flows with embedded shock waves can be calculated. To facilitate airfoil design, the method allows alternating between inverse and direct calculations to obtain a profile shape that satisfies given geometric constraints. Examples are shown of the application of the technique to improve the performance of several lifting airfoil sections. The extension of the method to three dimensions for designing supercritical wings is also indicated.
A review of hot wire anemometry in transonic flows
Stainback, P. C.
1985-01-01
The present paper provides a review of hot wire anemometry for compressible flows, giving particular attention to the transonic flow problem. It is pointed out that the first and most important definitive work in hot wire anemometry for compressible flows was reported by Kovasznay (1953). The existence of three independent fluctuating modes in compressible flows for small perturbations was found, taking into account the vorticity mode, the entropy mode, and the sound-wave mode. A review of Kovasznays' method for supersonic flows is also presented, and advances reported by Markovin (1956) are examined. Attention is given to experiments conducted by Horstman and Rose (1977), a general solution to the hot wire problem at transonic conditions sought by Stainback et al. (1983), and some apparent problems.
TRO-2D - A code for rational transonic aerodynamic optimization
Davis, W. H., Jr.
1985-01-01
Features and sample applications of the transonic rational optimization (TRO-2D) code are outlined. TRO-2D includes the airfoil analysis code FLO-36, the CONMIN optimization code and a rational approach to defining aero-function shapes for geometry modification. The program is part of an effort to develop an aerodynamically smart optimizer that will simplify and shorten the design process. The user has a selection of drag minimization and associated minimum lift, moment, and the pressure distribution, a choice among 14 resident aero-function shapes, and options on aerodynamic and geometric constraints. Design variables such as the angle of attack, leading edge radius and camber, shock strength and movement, supersonic pressure plateau control, etc., are discussed. The results of calculations of a reduced leading edge camber transonic airfoil and an airfoil with a natural laminar flow are provided, showing that only four design variables need be specified to obtain satisfactory results.
Basic numerical methods. [of unsteady and transonic flow
Steger, Joseph L.; Van Dalsem, William R.
1989-01-01
Some of the basic finite-difference schemes that can be used to solve the nonlinear equations that describe unsteady inviscid and viscous transonic flow are reviewed. Numerical schemes for solving the unsteady Euler and Navier-Stokes, boundary-layer, and nonlinear potential equations are described. Emphasis is given to the elementary ideas used in constructing various numerical procedures, not specific details of any one procedure.
Unsteady Interaction Between a Transonic Turbine Stage and Downstream Components
Davis Roger; Yao Jixian; Clark John; Stetson Gary; Alonso Juan; Jameson Antony; Haldeman Charles; Dunn Michael
2004-01-01
Results from a numerical simulation of the unsteady flow through one quarter of the circumference of a transonic high-pressure turbine stage, transition duct, and low-pressure turbine first vane are presented and compared with experimental data. Analysis of the unsteady pressure field resulting from the simulation shows the effects of not only the rotor/stator interaction of the high-pressure turbine stage but also new details of the interaction between the blade and the downstream transition...
Transonic analysis of complex configurations using TRANAIR program
Saaris, G. R.; Gilkey, R. D.; Smit, K. L.; Tinoco, E. N.
1989-01-01
The application of a three-dimensional transonic flow analysis method, TRANAIR, is explored from the point of view of a user. Detailed features of the program are outlined to give a better understanding of capability. Numerous results are presented to show some of the complex configurations which have been analyzed. In particular, examples are provided which show the application to turbofan engine installation on transport aircraft.
Distributed optical fiber perturbation sensing system based on Mach-Zehnder interferometer
Institute of Scientific and Technical Information of China (English)
Wengang WANG; Deming LIU; Hairong LIU; Qizhen SUN; Zhifeng SUN; Xu ZHANG; Ziheng XU
2009-01-01
A novel distributed optical fiber vibration-sensing system based on Mach-Zehnder interferometer has been designed and experimentally demonstrated. Firstly, the principle of Mach-Zehnder optical path interferometer technique is clarified. The output of the Mach-Zehnder interferometer is proportional to the phase shift induced by the perturbation. Secondly, the system consists of the laser diode (LD) as the light source, fiber, Mach-Zehnder optical interferometers as the sensing units, a 1×N star fiber-optic coupler, an N×1 fiber-optic coupler, a photodiode (PD) detector, and a computer used in signal processing. The entire monitoring region of this system is divided into several small zones, and each small monitoring zone is independent from each other. All of the small monitoring zones have their own sensing unit, which is defined by Mach-Zehnder optical interferometer. A series of sensing units are connected by the star fiber-optic couplers to form a whole sensing net. Thirdly, signal-processing techniques are subsequently used to calculate the phase shift to estimate whether intruders appear. The sensing system is able to locate the vibration signal simultaneously, includ-ing multiple vibrations at different positions, by employing the time-division multiplexed (TDM) technique. Finally, the operation performance of the proposed system is tested in the experiment lab with the conditions as follows: the number of the sensing units is 3, the length of the sensing fiber is 50 m, and the wavelength of the light diode is 1550nm. Based on these investigations, the fiber surrounding alert system is achieved. We have experimen-tally demonstrated that the sensing system can measure both the frequency and position of the vibration in real time, with a spatial positional resolution better than 50 m in an area of 1 km2.
Progress in the development of a Mach 5 quiet tunnel
Beckwith, I. E.; Andere, J. B.; Stainback, P. C.; Harvey, W. D.; Srokowski, A. J.
1977-01-01
Various techniques to control and reduce radiated noise and the application of these techniques to a 1/2-water Mach 5 quiet tunnel are reviewed. Measurements in a small scale nozzle have shown that the upstream part of the supersonic wall boundary layer could be maintained laminar up to Reynolds numbers of nearly 4 x 1 million based on the test region length upstream of the nozzle exit. Turbulent noise levels in this test region were then reduced by an order of magnitude. To maintain low noise levels at higher Reynolds numbers, laminar flow noise shields are required. Data are presented for shields that consist of small diameter rods alined nearly parallel to the entrance flow with small gaps between the rods for boundary layer suction. Analysis and data presented on the noise shielding and reflection characteristics of flat plates and a rod-wall test panel indicate that freestream turbulent noise can be reduced by 70 to 90 deg at high Reynolds numbers. Performance estimates for the 1/2-meter tunnel are based on these results.
Design and construction of models for the National Transonic Facility, part 1
Young, C. P., Jr.
1985-01-01
The design and construction of models for the National Transonic Facility (NTF) has resulted in significant technology developments in many areas. This lecture covers the development of design criteria and major research and development work that has contributed to the successful design and fabrication models for testing at full scale Reynolds number the NTF. Emphasis is placed on the materials aspect of the design and fabrication proces, including metallic materials, mechanical properties characterization, new steel alloy development, fracture toughness enhancement, and identification of fillers and solders suitable for use in cryogenic models. Quantitative data are provided which will be of value to the potential user of NTF or for application to the design and fabrication of model systems for other cryogenic wind tunnels.
Design and construction of models for the National Transonic Facility, part 2
Young, C. P., Jr.
1985-01-01
This lecture presents the results of fastener load and retention systems tests which were carried out as a part of the cryogenic models technology development program a the NASA Langley Research Center (LaRC). Various design concepts for the National Transonic Facility (NTF) developmental and production models are discussed. A number of NTF models are described with emphasis on materials used, uniqueness of design and design drivers. Design and fabrication experience is presented in terms of the primary thermal and mechanical considerations required for design as well as fabrication. Cost considerations are addressed in terms of factors influencing costs for NTF models and cost data comparisons which are taken from both NASA Langley and industry experience.
CFD Simulation of Transonic Flow in High-Voltage Circuit Breaker
Directory of Open Access Journals (Sweden)
Xiangyang Ye
2012-01-01
Full Text Available A high-voltage circuit breaker is an indispensable piece of equipment in the electric transmission and distribution systems. Transonic flow typically occurs inside breaking chamber during the current interruption, which determines the insulating characteristics of gas. Therefore, accurate compressible flow simulations are required to improve the prediction of the breakdown voltages in various test duties of high-voltage circuit breakers. In this work, investigation of the impact of the solvers on the prediction capability of the breakdown voltages in capacitive switching is presented. For this purpose, a number of compressible nozzle flow validation cases have been presented. The investigation is then further extended for a real high-voltage circuit breaker geometry. The correlation between the flow prediction accuracy and the breakdown voltage prediction capability is identified.
Blade Parameterization and Aerodynamic Design Optimization for a 3D Transonic Compressor Rotor
Institute of Scientific and Technical Information of China (English)
Naixing Chen; Hongwu Zhang; Yanji Xu; Weiguang Huang
2007-01-01
The present paper describes an optimization methodology for aerodynamic design of turbomachinery combined with a rapid 3D blade and grid generator (RAPID3DGRID), a N.S. solver, a blade parameterization method (BPM), a gradient-based parameterization-analyzing method (GPAM), a response surface method (RSM) with zooming algorithm and a simple gradient method. By the use of blade parameterization method a transonic compressor rotor can be expressed by a set of polynomials, and then it enables us to transform coordinate-expressed blade data to parameter-expressed and then to reduce the number of parameters. With changing any one of the parameters and by applying grid generator and N.S. solver, we can obtain several groups of samples. Here only ten parameters were considered to search an optimized compressor rotor. As a result of optimization, the adiabatic efficiency was increased by 1.73%.
Design and construction of 2 transonic airfoil models for tests in the NASA Langley C.3-M TCT
Schaechterle, G.; Ludewig, K. H.; Stanewsky, E.; Ray, E. J.
1982-01-01
As part of a NASA/DFVLR cooperation program two transonic airfoils were tested in the NASA Langley 0.3-m TCT. Model design and construction was carried out by DFVLR. The models designed and constructed performed extremely well under cryogenic conditions. Essentially no permanent changes in surface quality and geometric dimensions occurred during the tests. The aerodynamic results from the TCT tests which demonstrate the large sensitivity of the airfoil CAST 10-Z/DOAZ to Reynolds number changes compared well with results from other facilities at ambient temperatures.
Nonlinear dynamics approach of modeling the bifurcation for aircraft wing flutter in transonic speed
DEFF Research Database (Denmark)
Matsushita, Hiroshi; Miyata, T.; Christiansen, Lasse Engbo
2002-01-01
The procedure of obtaining the two-degrees-of-freedom, finite dimensional. nonlinear mathematical model. which models the nonlinear features of aircraft flutter in transonic speed is reported. The model enables to explain every feature of the transonic flutter data of the wind tunnel tests...
Mach band type lateral inhibition in different sense organs.
von Békésy, G
1967-01-01
Experiments were done on the skin with shearing forces, vibrations, and heat stimuli and on the tongue with taste stimuli to show that the well known Mach bands are not exclusively a visual phenomenon. On the contrary, it is not difficult to produce areas of a decreased sensation magnitude corresponding to the dark Mach bands in vision. It is shown on a geometrical model of nervous interaction that the appearance of Mach bands for certain patterns of stimulus distribution is correlated with nervous inhibition surrounding the area of sensation. This corroborates the earlier finding that surrounding every area transmitting sensation there is an area simultaneously transmitting inhibition.
The Influence of Ernst Mach in the Teaching of Mechanics
Assis, Andre K. T.; Zylbersztajn, Arden
We present Newton's main ideas for the formulation of classical mechanics as given in the Principia. Then we discuss Ernst Mach's criticisms of Newtonian mechanics as contained in his book The Science of Mechanics. We analyze the influence of Mach's ideas in the teaching of classical mechanics considering five representative textbooks: those of Kittel, Knight and Ruderman; Marion and Thornton; Symon; Feynman, Leighton and Sands; and Goldstein. We conclude that the influence of Mach's ideas has been very great, being incorporated in the textbooks, although not always with the deserved acknowledgment.
Mach-Zehnder Phasing Sensor for Elts
Dohlen, Kjetil; Montoya-Martinez, Luzma
Segmented mirror technology has been successfully applied to 10m class telescopes (Keck HET GTC) and its application to future extremely large telescopes (20m NG-CFHT 30m CELT 50m EURO50 100m OWL) is required. Extensive use of adaptive optics in these telescopes puts stringent specifications on wavefront error allowing typically of the order of lambda/20 to segmentation errors. Several phasing metrology schemes adaptable to these giant telescopes are under development. We investigate a novel technique based on the Mach-Zehnder interferometer with a spatial filter in one arm. Atmospheric turbulence is tolerated in this setup if the spatial filter has the size similar to that of the seeing disk. The resulting interference pattern only contains the high-frequency spatial information including information about the piston step height. We describe the theoretical analysis of this system and show simulated and experimatal results. Different error sources are analyzed in order to provide a preliminary idea of the merits of this technique compared with other phasing techniques.
Mach 6 flowfield survey at the engine inlet of a research airplane
Johnson, C. B.; Lawing, P. L.
1977-01-01
A flowfield survey was conducted to better define the nature of vehicle forebody flowfield at the inlet location of an airframe-integrated scramjet engine mounted on the lower surface of a high-speed research airplane to be air launched from a B-52 and rocket boosted to Mach 6. The tests were conducted on a 1/30-scale brass model in a Mach-6 20-in. wind tunnel at Reynolds number of 11,200,000 based on distance to engine inlet. Boundary layer profiles at five spanwise locations indicate that the boundary layer in the area of the forebody centerline is more than twice as thick as the boundary layer at three outboard stations. It is shown that the cold streak found in heating contours on the centerline of the forebody is caused by a thickening of the boundary layer on the centerline, and that this thickening decreases with angle of attack.
A fast spatial scanning combination emissive and mach probe for edge plasma diagnosis
Energy Technology Data Exchange (ETDEWEB)
Lehmer, R.D.; LaBombard, B.; Conn, R.W.
1989-04-01
A fast spatially scanning emissive and mach probe has been developed for the measurement of plasma profiles in the PISCES facility at UCLA. A pneumatic cylinder is used to drive a multiple tip probe along a 15cm stroke in less than 400msec, giving single shot profiles while limiting power deposition to the probe. A differentially pumped sliding O-ring seal allows the probe to be moved between shots to infer two and three dimensional profiles. The probe system has been used to investigate the plasma potential, density, and parallel mach number profiles of the presheath induced by a wall surface and scrape-off-layer profile modifications in biased limiter simulation experiments. Details of the hardware, data acquisition electronics, and tests of probe reliability are discussed. 30 refs., 24 figs.
Krypton tagging velocimetry in a turbulent Mach 2.7 boundary layer
Zahradka, D.; Parziale, N. J.; Smith, M. S.; Marineau, E. C.
2016-05-01
The krypton tagging velocimetry (KTV) technique is applied to the turbulent boundary layer on the wall of the "Mach 3 Calibration Tunnel" at Arnold Engineering Development Complex (AEDC) White Oak. Profiles of velocity were measured with KTV and Pitot-pressure probes in the Mach 2.7 turbulent boundary layer comprised of 99 % {N}2/1 % Kr at momentum-thickness Reynolds numbers of {Re}_{\\varTheta }= 800, 1400, and 2400. Agreement between the KTV- and Pitot-derived velocity profiles is excellent. The KTV and Pitot velocity data follow the law of the wall in the logarithmic region with application of the Van Driest I transformation. The velocity data are analyzed in the outer region of the boundary layer with the law of the wake and a velocity-defect law. KTV-derived streamwise velocity fluctuation measurements are reported and are consistent with data from the literature. To enable near-wall measurement with KTV (y/δ ≈ 0.1-0.2), an 800-nm longpass filter was used to block the 760.2-nm read-laser pulse. With the longpass filter, the 819.0-nm emission from the re-excited Kr can be imaged to track the displacement of the metastable tracer without imaging the reflection and scatter from the read-laser off of solid surfaces. To operate the Mach 3 AEDC Calibration Tunnel at several discrete unit Reynolds numbers, a modification was required and is described herein.
Improvement of Flow Quality in NAL Chofu Mach 10 Nozzle
Lacey, John; Inoue, Yasutoshi; Higashida, Akio; Inoue, Manabu; Ishizaka, Kouichi; Korte, John J.
2002-01-01
As a result of CFD analysis and remachining of the nozzle, the flow quality of the Mach 10 Hypersonic Wind Tunnel at NAL Chofu, Japan was improved. The subsequent test results validated the CFD analytical predictions by NASA and MHL.
Mach's Principle selects 4 space-time dimensions
Altshuler, Boris L
2012-01-01
Bi-tensor kernel in integral form of Einstein equations realizing Mach's idea of non-existence of empty space-times is taken as an inverse of differential operator ("Mach operator") defined conventionally as a second variation of Einstein's gravity Action over contravariant components of metric tensor. The choice of transverse gauge condition used in this definition does not influence results of the paper since only transverse and traceless tensor modes written on different background space-times are studied. Presence of ghosts among modes of Mach operator invalidates the integral formulation of Einstein equations. And the demand of absence of these ghosts proves to be a selection rule for dimensionality of the background space-time. In particular Mach operator written on De Sitter background or on the background of so called "Einstein Universe" does not possess tensor ghosts only in 4-dimensions. The similar demand gives non-trivial formula for dimensionalities of subspaces of the Freund-Rubin background.
Mach-Zehnder Fiber-Optic Links for ICF Diagnostics
Energy Technology Data Exchange (ETDEWEB)
Miller, E. K., Hermann, H. W.
2012-11-01
This article describes the operation and evolution of Mach-Zehnder links for single-point detectors in inertial confinement fusion experimental facilities, based on the Gamma Reaction History (GRH) diagnostic at the National Ignition Facility.
Improvement of Flow Quality in NAL Chofu Mach 10 Nozzle
Lacey, John; Inoue, Yasutoshi; Higashida, Akio; Inoue, Manabu; Ishizaka, Kouichi; Korte, John J.
2002-01-01
As a result of CFD analysis and remachining of the nozzle, the flow quality of the Mach 10 Hypersonic Wind Tunnel at NAL Chofu, Japan was improved. The subsequent test results validated the CFD analytical predictions by NASA and MHL.
Hryniewicki, M. K.; Gottlieb, J. J.; Groth, C. P. T.
2017-07-01
The transition boundary separating the region of regular reflection from the regions of single-, transitional-, and double-Mach reflections for a planar shock wave moving in air and interacting with an inclined wedge in a shock tube is studied by both analytical methods and computational-fluid-dynamic simulations. The analytical solution for regular reflection and the corresponding solutions from the extreme-angle (detachment), sonic, and mechanical-equilibrium transition criteria by von Neumann (Oblique reflection of shocks, Explosive Research Report No. 12, Navy Department, Bureau of Ordnance, U.S. Dept. Comm. Tech. Serv. No. PB37079 (1943). Also, John von Neumann, Collected Works, Pergamon Press 6, 238-299, 1963) are first revisited and revised. The boundary between regular and Mach reflection is then determined numerically using an advanced computational-fluid-dynamics algorithm to solve Euler's inviscid equations for unsteady motion in two spatial dimensions. This numerical transition boundary is determined by post-processing many closely stationed flow-field simulations, to determine the transition point when the Mach stem of the Mach-reflection pattern just disappears and this pattern then transcends into that of regular reflection. The new numerical transition boundary is shown to agree well with von Neumann's closely spaced sonic and extreme-angle boundaries for weak incident shock Mach numbers from 1.0 to 1.6, but this new boundary trends upward and above von Neumann's sonic and extreme-angle boundaries by a couple of degrees at larger shock Mach numbers from 1.6 to 4.0. Furthermore, the new numerically determined transition boundary is shown to agree well with very few available experimental data obtained from previous experiments designed to reflect two symmetrical moving oblique shock waves along a plane without a shear or boundary layer.
Hryniewicki, M. K.; Gottlieb, J. J.; Groth, C. P. T.
2016-12-01
The transition boundary separating the region of regular reflection from the regions of single-, transitional-, and double-Mach reflections for a planar shock wave moving in air and interacting with an inclined wedge in a shock tube is studied by both analytical methods and computational-fluid-dynamic simulations. The analytical solution for regular reflection and the corresponding solutions from the extreme-angle (detachment), sonic, and mechanical-equilibrium transition criteria by von Neumann (Oblique reflection of shocks, Explosive Research Report No. 12, Navy Department, Bureau of Ordnance, U.S. Dept. Comm. Tech. Serv. No. PB37079 (1943). Also, John von Neumann, Collected Works, Pergamon Press 6, 238-299, 1963) are first revisited and revised. The boundary between regular and Mach reflection is then determined numerically using an advanced computational-fluid-dynamics algorithm to solve Euler's inviscid equations for unsteady motion in two spatial dimensions. This numerical transition boundary is determined by post-processing many closely stationed flow-field simulations, to determine the transition point when the Mach stem of the Mach-reflection pattern just disappears and this pattern then transcends into that of regular reflection. The new numerical transition boundary is shown to agree well with von Neumann's closely spaced sonic and extreme-angle boundaries for weak incident shock Mach numbers from 1.0 to 1.6, but this new boundary trends upward and above von Neumann's sonic and extreme-angle boundaries by a couple of degrees at larger shock Mach numbers from 1.6 to 4.0. Furthermore, the new numerically determined transition boundary is shown to agree well with very few available experimental data obtained from previous experiments designed to reflect two symmetrical moving oblique shock waves along a plane without a shear or boundary layer.
Experimental Investigation of 'Transonic Resonance' with Convergent-Divergent Nozzles
Zaman, K. B. M. Q.; Dahl, M. D.; Bencic, T. J.; Zaman, Khairul (Technical Monitor)
2001-01-01
Convergent-divergent nozzles, when run at pressure ratios lower than the design value, often undergo a flow resonance accompanied by the emission of acoustic tones. The phenomenon, different in characteristics from conventional 'screech' tones, has been studied experimentally. Unlike screech, the frequency increases with increasing supply pressure. There is a 'staging' behavior; 'odd harmonic' stages resonate at lower pressures while the fundamental occurs in a range of higher pressures corresponding to a fully expanded Mach number (M(sub j)) around unity. The frequency (f(sub N)) variation with M(sub j) depends on the half angle-of-divergence (theta) of the nozzle. At smaller theta, the slope of f(sub N) versus M(sub j) curve becomes steeper. The resonance involves standing waves and is driven by unsteady shock/boundary layer interaction. The distance between the foot of the shock and the nozzle exit imposes the lengthscale (L'). The fundamental corresponds to a quarterwave resonance, the next stage at a lower supply pressure corresponds to a three-quarter-wave resonance, and so on. The principal trends in the frequency variation are explained simply from the characteristic variation of the length-scale L'. Based on the data, correlation equations are provided for the prediction of f(sub N). A striking feature is that tripping of the boundary layer near the nozzle's throat tends to suppress the resonance. In a practical nozzle a tendency for the occurrence of the phenomenon is thought to be a source of 'internal noise'; thus, there is a potential for noise benefit simply by appropriate boundary layer tripping near the nozzle's throat.
Igarashi, Asuka; Nitta, Shin-ya
2014-01-01
We have classified possible transonic solutions of galactic outflows in the gravitational potential of the dark matter halo (DMH) and super massive black hole (SMBH) under the assumptions of isothermal, spherically symmetric and steady state. It is clarified that the gravity of SMBH adds a new branch of transonic solutions with the transonic point in very close proximity to the centre in addition to the outer transonic point generated by the gravity of DMH. Because these two transonic solutions have substantially different mass fluxes and starting points, these solutions may have different influences on the evolution of galaxies and the release of metals into intergalactic space. We have applied our model to the Sombrero galaxy and obtained a new type of galactic outflow: a slowly accelerated transonic outflow through the transonic point at very distant region ($\\simeq 126$\\ kpc). In this galaxy, previous works reported a discrepancy that although the trace of the galactic outflow is observed by X-ray, the ga...
Numerical simulations of Mach stem formation via intersecting bow shocks
Hansen, E. C.; Frank, A.; Hartigan, P.; Yirak, K.
2015-12-01
Hubble Space Telescope observations show bright knots of Hα emission within outflowing young stellar jets. Velocity variations in the flow create secondary bow shocks that may intersect and lead to enhanced emission. When the bow shocks intersect at or above a certain critical angle, a planar shock called a Mach stem is formed. These shocks could produce brighter Hα emission since the incoming flow to the Mach stem is parallel to the shock normal. In this paper we report first results of a study using 2-D numerical simulations designed to explore Mach stem formation at the intersection of bow shocks formed by hypersonic "bullets" or "clumps". Our 2-D simulations show how the bow shock shapes and intersection angles change as the adiabatic index γ changes. We show that the formation or lack of a Mach stem in our simulations is consistent with the steady-state Mach stem formation theory. Our ultimate goal, which is part of an ongoing research effort, is to characterize the physical and observational consequences of bow shock intersections including the formation of Mach stems.
3D shock-bubble interactions at Mach 3
Hejazialhosseini, Babak; Koumoutsakos, Petros
2012-01-01
We present a simulation for the interactions of shockwaves with light spherical density inhomogeneities. Euler equations for two-phase compressible flows are solved in a 3D uniform resolution finite volume based solver using 5th order WENO reconstructions of the primitive quantities, HLL-type numerical fluxes and 3rd order TVD time stepping scheme. In this study, a normal Mach 3 shockwave in air is directed at a helium bubble with an interface Atwood number of -0.76. We employ 4 billion cells on a supercomputing cluster and demonstrate the development of this flow until relatively late times. Shock passage compresses the bubble and deposits baroclinic vorticity on the interface. Initial distribution of the vorticity and compressions lead to the formation of an air jet, interface roll-ups and the formation of a long lasting vortical core, the white core. Compressed upstream of the bubble turns into a mixing zone and as the vortex ring distances from this mixing zone, a plume-shaped region is formed and sustain...
Integrated Mach-Zehnder interferometer for Bose-Einstein condensates.
Berrada, T; van Frank, S; Bücker, R; Schumm, T; Schaff, J-F; Schmiedmayer, J
2013-01-01
Particle-wave duality enables the construction of interferometers for matter waves, which complement optical interferometers in precision measurement devices. This requires the development of atom-optics analogues to beam splitters, phase shifters and recombiners. Integrating these elements into a single device has been a long-standing goal. Here we demonstrate a full Mach-Zehnder sequence with trapped Bose-Einstein condensates confined on an atom chip. Particle interactions in our Bose-Einstein condensate matter waves lead to a nonlinearity, absent in photon optics. We exploit it to generate a non-classical state having reduced number fluctuations inside the interferometer. Making use of spatially separated wave packets, a controlled phase shift is applied and read out by a non-adiabatic matter-wave recombiner. We demonstrate coherence times a factor of three beyond what is expected for coherent states, highlighting the potential of entanglement as a resource for metrology. Our results pave the way for integrated quantum-enhanced matter-wave sensors.
Optimum Transonic Airfoils Based on the Euler Equations
Iollo, Angelo; Salas, Manuel, D.
1996-01-01
We solve the problem of determining airfoils that approximate, in a least square sense, given surface pressure distributions in transonic flight regimes. The flow is modeled by means of the Euler equations and the solution procedure is an adjoint- based minimization algorithm that makes use of the inverse Theodorsen transform in order to parameterize the airfoil. Fast convergence to the optimal solution is obtained by means of the pseudo-time method. Results are obtained using three different pressure distributions for several free stream conditions. The airfoils obtained have given a trailing edge angle.
An inverse method with regularity condition for transonic airfoil design
Zhu, Ziqiang; Xia, Zhixun; Wu, Liyi
1991-01-01
It is known from Lighthill's exact solution of the incompressible inverse problem that in the inverse design problem, the surface pressure distribution and the free stream speed cannot both be prescribed independently. This implies the existence of a constraint on the prescribed pressure distribution. The same constraint exists at compressible speeds. Presented here is an inverse design method for transonic airfoils. In this method, the target pressure distribution contains a free parameter that is adjusted during the computation to satisfy the regularity condition. Some design results are presented in order to demonstrate the capabilities of the method.
Aerodynamic Optimum Design of Transonic Turbine Cascades Using Genetic Algorithms
Institute of Scientific and Technical Information of China (English)
无
1997-01-01
This paper presents an aerodynamic optimum design method for transonic turbine cascades based on the Genetic Algorithms coupled to the inviscid flow Euler Solver and the boundary-layer calculation.The Genetic Algorithms control the evolution of a population of cascades towards an optimum design.The fitness value of each string is evaluated using the flow solver.The design procedure has been developed and the behavior of the genetic algorithms has been tested.The objective functions of the design examples are the minimum mean-square deviation between the aimed pressure and computed pressure and the minimum amount of user expertise.
Aeroelastic Tailoring of Transport Wings Including Transonic Flutter Constraints
Stanford, Bret K.; Wieseman, Carol D.; Jutte, Christine V.
2015-01-01
Several minimum-mass optimization problems are solved to evaluate the effectiveness of a variety of novel tailoring schemes for subsonic transport wings. Aeroelastic stress and panel buckling constraints are imposed across several trimmed static maneuver loads, in addition to a transonic flutter margin constraint, captured with aerodynamic influence coefficient-based tools. Tailoring with metallic thickness variations, functionally graded materials, balanced or unbalanced composite laminates, curvilinear tow steering, and distributed trailing edge control effectors are all found to provide reductions in structural wing mass with varying degrees of success. The question as to whether this wing mass reduction will offset the increased manufacturing cost is left unresolved for each case.
Simulated transonic flows for aircraft with nacelles, pylons, and winglets
Boppe, C. W.; Stern, M. A.
1980-01-01
A computational method which simulates transonic flow about wing-fuselage configurations has been extended to include the treatment of multiple body and non-planar wing surfaces. The finite difference relaxation scheme is characterized by a modified small disturbance flow equation and multiple embedded grid system. Wing-body combinations with as many as four nacelles/pods, four pylons, and wing-tip-mounted winglets can be analyzed. A scheme for modeling inlet spillage and engine exhaust interference effects has been included. Computed results are correlated with experimental data for three transport configurations.
Numerical evaluation of tandem rotor for highly loaded transonic fan
Institute of Scientific and Technical Information of China (English)
ZHAO Bin; LIU Bao-jie
2011-01-01
Transonic tandem rotor was designed for highly loaded fan at a corrected tip speed of 381 m/s and another conventional rotor was designed as a baseline to evaluate the loading superiority of tandem rotor with three-dimensional (3-D) numerical simulation. The aft blade solidity and its impact on total loading level were studied in depth. The result indicates that tandem rotor has potential to achieve higher loading level and attain favorable aerodynamic performance in a wide range of loading coefficient 0. 55 ～ 0.68, comparing with the conventional rotor which produced a total pressure ratio of 2.0 and loading coefficient of 0. 42.
Wind-US Unstructured Flow Solutions for a Transonic Diffuser
Mohler, Stanley R., Jr.
2005-01-01
The Wind-US Computational Fluid Dynamics flow solver computed flow solutions for a transonic diffusing duct. The calculations used an unstructured (hexahedral) grid. The Spalart-Allmaras turbulence model was used. Static pressures along the upper and lower wall agreed well with experiment, as did velocity profiles. The effect of the smoothing input parameters on convergence and solution accuracy was investigated. The meaning and proper use of these parameters are discussed for the benefit of Wind-US users. Finally, the unstructured solver is compared to the structured solver in terms of run times and solution accuracy.
Abramova, K. A.; Petrov, A. V.; Potapchick, A. V.; Soudakov, V. G.
2016-10-01
Numerical and experimental investigations of transonic buffet control by tangential jet blowing are presented. To suppress the shock-induced boundary layer separation and the buffet at transonic speeds, compressed air jet is blown through a small slot nozzle tangentially to the upper surface of the supercritical airfoil. Numerical simulations were carried out on the basis of the unsteady Reynolds averaged Navier-Stokes (URANS) equations. Experimental studies of the tangential jet blowing were performed in the transonic wind tunnel T-112 of TsAGI. Results show that the jet moves the shock downstream, increases lift, suppresses flow separation under shock foot and delays buffet onset.
Quasi-normal acoustic oscillations in the transonic Bondi flow
Chaverra, Eliana
2015-01-01
In recent work, we analyzed the dynamics of spherical and nonspherical acoustic perturbations of the Michel flow, describing the steady radial accretion of a relativistic perfect fluid into a nonrotating black hole. We showed that such perturbations undergo quasi-normal oscillations and computed the corresponding complex frequencies as a function of the black hole mass M and the radius r_c of the sonic horizon. It was found that when r_c is much larger than the Schwarzschild radius r_H = 2GM/c^2 of the black hole, these frequencies scale like the surface gravity of the analogue black hole associated with the acoustic metric. In this work, we analyze the Newtonian limit of the Michel solution and its acoustic perturbations. In this limit, the flow outside the sonic horizon reduces to the transonic Bondi flow, and the acoustic metric reduces to the one introduced by Unruh in the context of experimental black hole evaporation. We show that for the transonic Bondi flow, Unruh's acoustic metric describes an analog...
Two-Fluid Equilibrium for Transonic Poloidal Flows
Guazzotto, Luca; Betti, Riccardo
2012-03-01
Much analytical and numerical work has been done in the past on ideal MHD equilibrium in the presence of macroscopic flow. In recent years, several authors have worked on equilibrium formulations for a two-fluid system, in which inertial ions and massless electrons are treated as distinct fluids. In this work, we present our approach to the formulation of the two-fluid equilibrium problem. Particular attention is given to the relation between the two-fluid equations and the equilibrium equations for the single-fluid ideal MHD system. Our purpose is to reconsider the results of one-fluid calculation with the more accurate two-fluid model, referring in particular to the so-called transonic discontinuities, which occur when the poloidal velocity spans a range crossing the poloidal sound speed (i.e., the sound speed reduced by a factor Bp/B). It is expected that the one-fluid discontinuity will be resolved into a sharp gradient region by the two-fluid model. Also, contrary to the ideal MHD case, in the two-fluid model the equations governing the equilibrium are elliptic in the whole range of interest for transonic equilibria. The numerical solution of the two-fluid system of equations is going to be based on a code built on the structure of the existing ideal-MHD code FLOW.
Numerical Simulations of Mach Stem Formation via Intersecting Bow Shocks
Hansen, Edward C; Hartigan, Patrick
2014-01-01
Hubble Space Telescope observations show bright knots of H$\\alpha$ emission within outflowing young stellar jets. Velocity variations in the flow create secondary bow shocks that may intersect and lead to enhanced emission. When the bow shocks intersect at or above a certain critical angle, a planar shock called a Mach stem is formed. These shocks could produce brighter H$\\alpha$ emission since the incoming flow to the Mach stem is parallel to the shock normal. In this paper we report first results of a study using 2-D numerical simulations designed to explore Mach stem formation at the intersection of bow shocks formed by hypersonic "bullets" or "clumps". Our 2-D simulations show how the bow shock shapes and intersection angles change as the adiabatic index $\\gamma$ changes. We show that the formation or lack of a Mach stem in our simulations is consistent with the steady-state Mach stem formation theory. Our ultimate goal, which is part of an ongoing research effort, is to characterize the physical and obse...
Mach-Zehnder recording systems for pulsed power diagnostics
Energy Technology Data Exchange (ETDEWEB)
Miller, E. K.; Abbott, R. Q.; McKenna, I.; Macrum, G.; Baker, D.; Tran, V.; Rodriguez, E.; Kaufman, M. I.; Tibbits, A.; Silbernagel, C. T.; Waltman, T. B. [National Security Technologies, LLC, Santa Barbara and Livermore, California 93111 (United States); National Security Technologies, LLC, Los Alamos, New Mexico 87544 (United States); and National Security Technologies, LLC, North Las Vegas, Nevada 89193 (United States); Herrmann, H. W.; Kim, Y. H.; Mack, J. M.; Young, C. S.; Caldwell, S. E.; Evans, S. C.; Sedillo, T. J. [Los Alamos National Laboratory, Los Alamos, New Mexico 87545 (United States); Stoeffl, W.; Grafil, E. [Lawrence Livermore National Laboratory, Livermore, California (United States); and others
2012-10-15
Fiber-optic transmission and recording systems, based on Mach-Zehnder modulators, have been developed and installed at the National Ignition Facility (NIF), and are being developed for other pulsed-power facilities such as the Z accelerator at Sandia, with different requirements. We present the design and performance characteristics for the mature analog links, based on the system developed for the Gamma Reaction History diagnostic at the OMEGA laser and at NIF. For a single detector channel, two Mach-Zehnders are used to provide high dynamic range at the full recording bandwidth with no gaps in the coverage. We present laboratory and shot data to estimate upper limits on the radiation effects as they impact recorded data quality. Finally, we will assess the technology readiness level for mature and developing implementations of Mach-Zehnder links for these environments.
Mach-Zehnder recording systems for pulsed power diagnostics.
Miller, E K; Abbott, R Q; McKenna, I; Macrum, G; Baker, D; Tran, V; Rodriguez, E; Kaufman, M I; Tibbits, A; Silbernagel, C T; Waltman, T B; Herrmann, H W; Kim, Y H; Mack, J M; Young, C S; Caldwell, S E; Evans, S C; Sedillo, T J; Stoeffl, W; Grafil, E; Liebman, J; Beeman, B; Watts, P; Carpenter, A; Horsfied, C J; Rubery, M S; Chandler, G A; Torres, J A; Smelser, R M
2012-10-01
Fiber-optic transmission and recording systems, based on Mach-Zehnder modulators, have been developed and installed at the National Ignition Facility (NIF), and are being developed for other pulsed-power facilities such as the Z accelerator at Sandia, with different requirements. We present the design and performance characteristics for the mature analog links, based on the system developed for the Gamma Reaction History diagnostic at the OMEGA laser and at NIF. For a single detector channel, two Mach-Zehnders are used to provide high dynamic range at the full recording bandwidth with no gaps in the coverage. We present laboratory and shot data to estimate upper limits on the radiation effects as they impact recorded data quality. Finally, we will assess the technology readiness level for mature and developing implementations of Mach-Zehnder links for these environments.
Mach-Zehnder Recording Systems for Pulsed Power Diagnostics
Energy Technology Data Exchange (ETDEWEB)
Miller, E K; McKenna, I; Macrum, G; Baker, D; Tran, V; Rodriguez, E; Kaufman, M I; Tibbits, A; Silbernagel, C T; Waltman, T B; Herrmann, H W; Kim, Y H; Mack, J M; Young, C S; Caldwell, S E; Evans, S C; Sedillo, T J; Stoeffl, W; Grafil, E; Liebman, J; Beeman, B; Watts, P; Carpenter, A; Horsfied, C J; Rubery, M S; Chandler, G A; Torres, J A
2012-10-01
Fiber-optic transmission and recording systems, based on Mach-Zehnder modulators, have been developed and installed at the National Ignition Facility (NIF), and are being developed for other pulsed-power facilities such as Z-R at Sandia, with different requirements. We present the design and performance characteristics for the mature analog links, based on the system developed for the Gamma Reaction History (GRH) diagnostic at OMEGA and NIF. For a single detector channel, two Mach-Zehnders are used to provide high dynamic range at the full recording bandwidth with no gaps in the coverage. We present laboratory and shot data to estimate upper limits on the radiation effects as they impact recorded data quality. Finally, we will assess the technology readiness level for mature and developing implementations of Mach-Zehnder links for these environments.
Quantum heat engines based on electronic Mach-Zehnder interferometers
Hofer, Patrick P.; Sothmann, Björn
2015-05-01
We theoretically investigate the thermoelectric properties of heat engines based on Mach-Zehnder interferometers. The energy dependence of the transmission amplitudes in such setups arises from a difference in the interferometer arm lengths. Any thermoelectric response is thus of purely quantum-mechanical origin. In addition to an experimentally established three-terminal setup, we also consider a two-terminal geometry as well as a four-terminal setup consisting of two interferometers. We find that Mach-Zehnder interferometers can be used as powerful and efficient heat engines which perform well under realistic conditions.
Flegel, Ashlie B.; Giel, Paul W.; Welch, Gerard E.
2014-01-01
The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to -51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At
A Direct-Fire Trajectory Model for Supersonic, Transonic, and Subsonic Projectile Flight
2014-07-01
motions of the projectile about the trajectory due to the angular motion of the projectile . For a stable projectile , these motions are typically small...A Direct-Fire Trajectory Model for Supersonic, Transonic, and Subsonic Projectile Flight by Paul Weinacht ARL-TR-6998 July 2014...Direct-Fire Trajectory Model for Supersonic, Transonic, and Subsonic Projectile Flight Paul Weinacht Weapons and Materials Research Directorate, ARL
Computational Fluid Dynamic Model of Steam Ingestion into a Transonic Compressor
2009-06-01
DYNAMIC MODEL OF STEAM INGESTION INTO A TRANSONIC COMPRESSOR by Collin R. Hedges June 2009 Thesis Advisor: Anthony J. Gannon Second...TYPE AND DATES COVERED Master’s Thesis 4. TITLE AND SUBTITLE Computational Fluid Dynamic Model of Steam Ingestion into a Transonic Compressor 6...flight deck. When ingested into jet engines, this steam may increase the engines’ susceptibility to stall. The serpentine air inlet ducts and single
Transition to Double Mach Stem for Nuclear Explosion at 104 ft Height of Burst.
1981-11-17
intersecting the ground. The initialization provides a strong shock with Mach number MI = 12. This speed and the need for restart capability led to the choice...a HOB of 104 ft (31.7m). A strong spherical shock is created in the surrounding air, and’ reflects from the grcund. 9 The outward-traveling airbiast...AIR FCIPCF SYST T’M CCvfvtANC NORTON" A!7, CA 9?40Pg (MIIJ’r’-MAN) QICY ATTN "INNYH "D IALAN5S<Y 0O1C Y ATTNJ MMN)) eHM kF-LVECCHir OICY ATTN fuNN w
Study of Rayleigh scattering for visualization of helium-air mixing at Mach 6
Shirinzadeh, B.; Balla, R. J.; Hillard, M. E.; Anders, J. B.; Exton, R. J.; Waitz, I. A.
1991-01-01
Using an ArF excimer laser, planar Rayleigh scattering measurements were performed to investigate helium mixing into air at supersonic speeds. These experiments were conducted in the Mach 6, high-Reynolds-number facility at NASA Langley Research Center. The capability of the Rayleigh scattering technique for flow visualization of a turbulent environment was demonstrated. The qualitative agreement between the averaged Rayleigh results and the reduced mean-mass-densities obtained from probe measurements substantiate that careful application of the technique, even in the presence of clusters, can give very useful results. It was also demonstrated that planar, quantitative measurements can be made in the absence of clusters.
Experimental Study on the Dynamic Stability of the IXV Configuration
Gülhan, Ali; Klevanski, Josef; Gawehn, Thomas
2015-01-01
Dynamic stability of the IXV configuration has been investigated using free oscillation measurement technique in the Trisonic Windtunnel (TMK). In the transonic Mach number range an escalating behavior of the pitching moment damping derivative has been observed, although the vehicle is statically stable. At Mach 0.8 the vehicle showed the most unstable behavior. The instability becomes weaker with increasing Mach number. At Mach number 1.1 the vehicle is only slight...
Energy Technology Data Exchange (ETDEWEB)
Huerst, C.; Schulz, A.; Wittig, S. [Univ. Karlsruhe (Germany). Lehrstuhl und Inst. fuer Thermische Stroemungsmaschinen
1995-04-01
The present study compares measured and computed heat transfer coefficients for high-speed boundary layer nozzle flows under engine Reynolds number conditions (U{sub {infinity}} = 230 {divided_by} 880 m/s, Re* = 0.37 {divided_by} 1.07 {times} 10{sup 6}). Experimental data have been obtained by heat transfer measurements in a two-dimensional, nonsymmetric, convergent-divergent nozzle. The nozzle wall is convectively cooled using water passages. The coolant heat transfer data and nozzle surface temperatures are used as boundary conditions for a three-dimensional finite-element code, which is employed to calculate the temperature distribution inside the nozzle wall. Heat transfer coefficients along the hot gas nozzle wall are derived from the temperature gradients normal to the surface. The results are compared with numerical heat transfer predictions using the low-Reynolds-number {kappa}-{epsilon} turbulence model by Lam and Bremhorst. Influence of compressibility in the transport equations for the turbulence properties is taken into account by using the local averaged density. The results confirm that this simplification leads to good results for transonic and low supersonic flows.
Effects of the composition on transonic properties of accretion flows around black holes
Chattopadhyay, Indranil
2013-01-01
We study the properties of a steady, multi-species, low angular momentum accretion flow around a Schwarzschild black hole. Each species is described by a relativistic equation of state. We find that the transonic properties depend strongly on the composition of the flow. We find that an electron-positron pair plasma is the least relativistic one. This flow produces only one sonic point very close to the event horizon and does not show multiple critical points for any angular momentum or energy. When the baryons are present, the number of critical points depend on the specific energy content. Since the number of critical points decide whether the flow will have non-linearities or shock waves, our results implies that whether or not standing shocks forms will depend on the flow composition. Thus, for instance, a pure electron-positron pair plasma will never undergo a shock transition, while mixing it with some baryons (common in outflows and jets, for example) as in a completely ionized gas, will have shocks. W...
AERODYNAMIC OPTIMIZATION DESIGN OF LOW ASPECT RATIO TRANSONIC TURBINE STAGE
Institute of Scientific and Technical Information of China (English)
SONG Liming; LI Jun; FENG Zhenping
2006-01-01
The advanced optimization method named as adaptive range differential evolution (ARDE)is developed. The optimization performance of ARDE is demonstrated using a typical mathematical test and compared with the standard genetic algorithm and differential evolution. Combined with parallel ARDE, surface modeling method and Navier-Stokes solution, a new automatic aerodynamic optimization method is presented. A low aspect ratio transonic turbine stage is optimized for the maximization of the isentropic efficiency with forty-one design variables in total. The coarse-grained parallel strategy is applied to accelerate the design process using 15 CPUs. The isentropic efficiency of the optimum design is 1.6% higher than that of the reference design. The aerodynamic performance of the optimal design is much better than that of the reference design.
The efficient solution of transonic wing flow fields
Holst, T. L.; Subramanian, N. R.; Thomas, S. D.
1983-01-01
An evaluation of the transonic-wing-analysis computer code TWING is presented. TWING utilizes a fully implicit, approximate-factorization iteration scheme to solve the full-potential equation in conservative form. A numerical elliptic-solver grid-generation scheme is used to generate the required finite-difference mesh. Several wing configurations have been analyzed, and comparisons of computed results have been made with available experimental data. Results indicate that the code is robust, accurate (when significant viscous effects are not present), and efficient. TWING generally produces solutions an order of magnitude faster than other conservative, full-potential codes using successive-line overrelaxation. The present method is applicable to a wide range of isolated wing configurations, including high-aspect-ratio transport wings and low-aspect-ratio, high-sweep, fighter configurations.
Recent applications of the transonic wing analysis computer code, TWING
Subramanian, N. R.; Holst, T. L.; Thomas, S. D.
1982-01-01
An evaluation of the transonic-wing-analysis computer code TWING is given. TWING utilizes a fully implicit approximate factorization iteration scheme to solve the full potential equation in conservative form. A numerical elliptic-solver grid-generation scheme is used to generate the required finite-difference mesh. Several wing configurations were analyzed, and the limits of applicability of this code was evaluated. Comparisons of computed results were made with available experimental data. Results indicate that the code is robust, accurate (when significant viscous effects are not present), and efficient. TWING generally produces solutions an order of magnitude faster than other conservative full potential codes using successive-line overrelaxation. The present method is applicable to a wide range of isolated wing configurations including high-aspect-ratio transport wings and low-aspect-ratio, high-sweep, fighter configurations.
Recent Productivity Improvements to the National Transonic Facility
Popernack, Thomas G., Jr.; Sydnor, George H.
1998-01-01
Productivity gains have recently been made at the National Transonic Facility wind tunnel at NASA Langley Research Center. A team was assigned to assess and set productivity goals to achieve the desired operating cost and output of the facility. Simulations have been developed to show the sensitivity of selected process productivity improvements in critical areas to reduce overall test cycle times. The improvements consist of an expanded liquid nitrogen storage system, a new fan drive, a new tunnel vent stack heater, replacement of programmable logic controllers, an increased data communications speed, automated test sequencing, and a faster model changeout system. Where possible, quantifiable results of these improvements are presented. Results show that in most cases, improvements meet the productivity gains predicted by the simulations.
Relaxation techniques for three-dimensional transonic flow about wings.
Bailey, F. R.; Steger, J. L.
1972-01-01
A relaxation procedure has been developed to treat the three-dimensional, transonic small perturbation equations about finite lifting wings. A combination of two schemes is employed. For flow forward of the wing trailing edge the equations are written in terms of a velocity potential in order to minimize computer algebra and storage. For the remaining flow field the equations are written in terms of the velocity components in order to simplify the enforcement of the Kutta condition. Difference equations and relaxation procedures are described for both schemes. The computational method automatically captures the imbedded shock wave in the three-dimensional flow field. Computed results are given and compared to experiment and other inviscid methods.
Formation of multiple shocklets in a transonic diffuser flow
Handa, T.; Miyazato, Y.; Masuda, M.; Matsuo, K.
Multiple shocklets are frequently generated in transonic diffuser flows. The present paper investigates the formation of these shocklets with a high-speed CCD camera combined with the schlieren method. It is observed that compression waves steepen while propagating upstream, and eventually become new shock waves. The ordinary shock wave is found to move upstream beyond the nozzle throat or to disappear while moving downstream depending on the pressure ratio across the nozzle. This phenomenon is also analyzed with the one-dimensional Euler equations by assuming a pressure disturbance given by the sine function at the channel exit. The calculated results are found to reproduce quite well the experimental behavior of the shocklets. The effect of the frequency of disturbance is also studied numerically, and it is shown that the multiple shocklet pattern appears when the amplitude of disturbance is not large and the diverging part of the channel downstream of the ordinary shock wave is long.
Transonic Tones and Excess Broadband Noise in Overexpanded Supersonic Jets
Zaman, Khairul B. M. Q.
2009-01-01
Noise characteristics of convergent-divergent (C-D) nozzles in the overexpanded regime are the focus of this paper. The flow regime is encountered during takeoff and landing of certain airplanes and also with rocket nozzles in launch-pad environment. Experimental results from laboratory-scale single nozzles are discussed. The flow often undergoes a resonance accompanied by emission of tones (referred to as transonic tones). The phenomenon is different from the well-known screech tones. Unlike screech, the frequency increases with increasing supply pressure. There is a staging behavior odd harmonic stages occur at lower pressures while the fundamental occurs in a range of relatively higher pressures. A striking feature is that tripping of the nozzle s internal boundary layer tends to suppress the resonance. However, even in the absence of tones the broadband levels are found to be high. That is, relative to a convergent case and at same pressure ratio, the C-D nozzles are found to be noisier, often by more than 10dB. This excess broadband noise (referred to as EBBN) is further explored. Its characteristics are found to be different from the well-known broadband shockassociated noise ( BBSN ). For example, while the frequency of the BBSN peak varies with observation angle no such variation is noted with EBBN. The mechanisms of the transonic tone and the EBBN are not completely understood yet. They appear to be due to unsteady shock motion inside the nozzle. The shock drives the flow downstream like a vibrating diaphragm, and resonance takes place similarly as with acoustic resonance of a conical section having one end closed and the other end open. When the boundary layer is tripped, apparently a breakdown of azimuthal coherence suppresses the resonance. However, there is still unsteady shock motion albeit with superimposed randomness. Such random motion of the internal shock and its interaction with the separated boundary layer produces the EBBN.
Flegel, Ashlie Brynn; Giel, Paul W.; Welch, Gerard E.
2014-01-01
The effects of inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The high turbulence study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Downstream total pressure and exit angle data were acquired for ten incidence angles ranging from +15.8 to 51.0. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12105 to 2.12106 and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 0.25 - 0.4 for the low Tu tests and 8- 15 for the high Tu study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitchyaw probe located in a survey plane 7 axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall
Concept Development of a Mach 1.6 High-Speed Civil Transport
Shields, Elwood W.; Fenbert, James W.; Ozoroski, Lori P.; Geiselhart, Karl A.
1999-01-01
A high-speed civil transport configuration with a Mach number of 1.6 was developed as part of the NASA High-Speed Research Program to serve as a baseline for assessing advanced technologies required for an aircraft with a service entry date of 2005. This configuration offered more favorable solutions to environmental concerns than configurations with higher Mach numbers. The Mach 1.6 configuration was designed for a 6500 n.mi. mission with a 250-passenger payload. The baseline configuration has a wing area of 8732 square feet a takeoff gross weight of 591570 lb, and four 41000-lb advanced turbine bypass engines defined by NASA. These engines have axisymmetric mixer-ejector nozzles that are assumed to yield 20 dB of noise suppression during takeoff, which is assumed to satisfy, the FAR Stage III noise requirements. Any substantial reduction in this assumed level of suppression would require oversizing the engines to meet community noise regulations and would severly impact the gross weight of the aircraft at takeoff. These engines yield a ratio of takeoff thrust to weight of 0.277 and a takeoff wing loading of 67.8 lb/square feet that results in a rotation speed of 169 knots. The approach velocity of the sized configuration at the end of the mission is 131 knots. The baseline configuration was resized with an engine having a projected life of 9000 hr for hot rotating parts and 18000 hr for the rest of the engine, as required for commercial use on an aircraft with a service entry date of 2005. Results show an increase in vehicle takeoff gross weight of approximately 58700 lb. This report presents the details of the configuration development, mass properties, aerodynamic design, propulsion system and integration, mission performance, and sizing.
Tunable multiwavelength erbium-doped fiber laser based on an in-line Mach Zehnder interferometer
Energy Technology Data Exchange (ETDEWEB)
Han, Young-Geun [Hanyang University, Seoul (Korea, Republic of)
2010-12-15
A tunable multiwavelength erbium-doped fiber laser based on an in-line Mach Zehnder interferometer is proposed and experimentally demonstrated. The in-line Mach Zehnder interferometer is realized by using cascaded long-period fiber gratings. The long-period fiber gratings can couple the guided core mode to several cladding modes. If two identical long-period fiber gratings are concatenated, an interference pattern can be generated, which results from an interaction of the core and the cladding modes in the second long-period fiber grating. Therefore, a simple multichannel filter based on an in-line Mach Zehnder interferometer can be realized. The wavelength spacing of the proposed multichannel filter is controlled by the number of long-period fiber gratings. We apply the proposed multichannel fiber to the generation of a multiwavelength erbium-doped fiber laser with a tunability on the order of the wavelength spacing. An erbium-doped fiber amplifier is implemented as a gain medium. The gain competition of erbium ions is suppressed by soaking the erbium-doped fiber in liquid nitrogen. The power fluctuation of the proposed multiwavelength fiber laser is measured to be less than 0.5 dB. A high-quality multiwavelength output with a high extinction ratio of more than 40 dB is achieved. The wavelength spacing of the proposed multiwavelength fiber laser is controlled by increasing the number of long-period fiber gratings. The wavelength spacing is changed from 0.8 nm to 1.6 nm discretely.
Interaction between a shock wave and a turbulent boundary layer in transonic flow
Adamson, T. C., Jr.; Feo, A.
1975-01-01
Interaction between a shock wave and an unseparated turbulent boundary layer is considered. The method of matched asymptotic expansions is used, with solutions valid in the double limit as Reynolds number tends to infinity and Mach number tends to unity. The shock is weak enough that interaction effects can be considered as perturbations to the undisturbed flow; the case considered is that where the sonic line is near the outer edge of the boundary layer. It is shown that, with order estimates for Reynolds stress perturbations, the induced wall pressure distribution can be calculated using only the two outer interaction regions, independent of a specific closure condition and that this solution is in fact a turbulent free interaction solution. A detailed analysis of the inner regions, for which an eddy viscosity model for the Reynolds shear stress is used, provides a description of the variations in velocity, temperature and density near and at the wall.
Investigation of a transonic separating/reattaching shear layer by means of PIV
Directory of Open Access Journals (Sweden)
S. Scharnowski
2015-01-01
Full Text Available The separating/reattaching flow over an axisymmetric backward-facing step is analyzed experimentally by means of particle image velocimetry (PIV. The main purpose of the measurements is the investigation of the mean flow field as well as of the Reynolds stress distributions at a Mach number of 0.7 and at a Reynolds number of 3.3×105 based on the step height. Due to the strong progress of optical flow measurements in the last years it was possible to resolve all flow scales down to 180μm (≈1% of the step height with high precision. Thanks to the high spatial resolution it was found for the first time that the Reynolds stress distribution features a local minimum between the first part of the shear layer and a region inside the recirculation region. This implies a more complex wake dynamics than assumed before.
Energy Technology Data Exchange (ETDEWEB)
Bozinoski, Radoslav; William Winters
2016-01-01
The purpose of this report is to document the theoretical models utilized by the computer code NETFLOW. This report will focus on the theoretical models used to analyze high Mach number fully compressible transonic flows in piping networks.
Weinstein, I.
1973-01-01
Heat-transfer and pressure distributions were measured over the surfaces of three hemisphere-cylinder models tested at a nominal Mach number of 7 in the Langley 8-foot high-temperature structures tunnel which uses methane-air products of combustion as a test medium. The results showed that the heat-transfer and pressure distributions over the surface of the models were in good agreement with experimental data obtained in air and also with theoretical predictions.
Cole, Stanley R.; Keller, Donald F.; Piatak, David J.
2000-01-01
The NASA Langley Transonic Dynamics Tunnel (TDT) has provided wind-tunnel experimental validation and research data for numerous launch vehicles and spacecraft throughout its forty year history. Most of these tests have dealt with some aspect of aeroelastic or unsteady-response testing, which is the primary purpose of the TDT facility. However, some space-related test programs that have not involved aeroelasticity have used the TDT to take advantage of specific characteristics of the wind-tunnel facility. In general. the heavy gas test medium, variable pressure, relatively high Reynolds number and large size of the TDT test section have made it the preferred facility for these tests. The space-related tests conducted in the TDT have been divided into five categories. These categories are ground wind loads, launch vehicle dynamics, atmospheric flight of space vehicles, atmospheric reentry. and planetary-probe testing. All known TDT tests of launch vehicles and spacecraft are discussed in this report. An attempt has been made to succinctly summarize each wind-tunnel test, or in the case of multiple. related tests, each wind-tunnel program. Most summaries include model program discussion, description of the physical wind-tunnel model, and some typical or significant test results. When available, references are presented to assist the reader in further pursuing information on the tests.
Effects of wind-tunnel noise on swept-cylinder transition at Mach 3.5
Creel, T. R., Jr.; Beckwith, I. E.; Chen, F.-J.
1986-01-01
Transition data are reported for circular cylinders at swept angles of 45 and 60 degrees in the Mach 3.5 pilot-low-disturbance tunnel where free-stream noise levels are varied from approximately .05-0.5 percent in terms of the rms fluctuating pressure normalized by the mean static pressure. Results indicate that end plate or boundary layer trip disturbances at the upstream end of the cylinders cause turbulent flow along the entire test Reynolds number range of 10-170 thousand per inch. With all end plate and trip disturbances removed, transition at the attachment lines occurred at free-stream Reynolds numbers based on diameters of about 70-80 thousand, independent of stream noise levels. The installation of small trips on the attachement lines caused transition at lower Reynolds numbers, depending on both the roughness height and the wind tunnel noise level.
3 TUNNELS IN THE ENGINE RESEARCH BUILDING ERB - IN CELL CE-26 VARIABLE REYNOLDS NUMBER SUPERSONIC NO
1956-01-01
3 TUNNELS IN THE ENGINE RESEARCH BUILDING ERB - IN CELL CE-26 VARIABLE REYNOLDS NUMBER SUPERSONIC NOZZLE - CELL CE-4 6X6 INCH MACH NUMBER 2.96 SUPERSONIC AIRPLANE - CELL 1-NW 1X1 FOOT MACH 3.12 SUPERSONIC TUNNEL
Mach-Zehnder fiber interferometer for people monitoring
Vasinek, Vladimir; Latal, Jan; Koudelka, Petr; Siska, Petr; Vitasek, Jan; Skapa, Jan
2010-10-01
Fiber optical interferometers belong to highly sensitive equipments that are able to measure slight changes like distortion of shape, temperature and electric field variation and etc. Their great advantage is that they are insensitive on ageing component, from which they are composed of. It is in virtue of herewith, that there are evaluated no changes in optical signal intensity but number interference fringes. To monitor the movement of persons, eventually to analyze the changes in state of motion we developed method based on analysis the dynamic changes in interferometric pattern. We have used Mach- Zehnder interferometer with conventional SM fibers excited with the DFB laser at wavelength of 1550 nm. It was terminated with optical receiver containing InGaAs PIN photodiode. Its output was brought into measuring card module that performs on FFT of the received interferometer signal. The signal rises with the composition of two waves passing through single interferometer arm. The optical fiber SMF 28e in one arm is referential; the second one is positioned on measuring slab at dimensions of 1x2m. A movement of persons over the slab was monitored, signal processed with FFT and frequency spectra were evaluated. They rose owing to dynamic changes of interferometric pattern. The results reflect that the individual subjects passing through slab embody characteristic frequency spectra, which are individual for particular persons. The scope of measuring frequencies proceeded from zero to 10 KHz. It was also displayed in experiments that the experimental subjects, who walked around the slab and at the same time they have had changed their state of motion (knee joint fixation), embodied characteristic changes in their frequency spectra. At experiments the stability of interferometric patterns was evaluated as from time aspects, so from the view of repeated identical experiments. Two kinds of balls (tennis and ping-pong) were used to plot the repeatability measurements and
Quantitative Global Heat Transfer in a Mach-6 Quiet Tunnel
Sullivan, John P.; Schneider, Steven P.; Liu, Tianshu; Rubal, Justin; Ward, Chris; Dussling, Joseph; Rice, Cody; Foley, Ryan; Cai, Zeimin; Wang, Bo; Woodiga, Sudesh
2012-01-01
This project developed quantitative methods for obtaining heat transfer from temperature sensitive paint (TSP) measurements in the Mach-6 quiet tunnel at Purdue, which is a Ludwieg tube with a downstream valve, moderately-short flow duration and low levels of heat transfer. Previous difficulties with inferring heat transfer from TSP in the Mach-6 quiet tunnel were traced to (1) the large transient heat transfer that occurs during the unusually long tunnel startup and shutdown, (2) the non-uniform thickness of the insulating coating, (3) inconsistencies and imperfections in the painting process and (4) the low levels of heat transfer observed on slender models at typical stagnation temperatures near 430K. Repeated measurements were conducted on 7 degree-half-angle sharp circular cones at zero angle of attack in order to evaluate the techniques, isolate the problems and identify solutions. An attempt at developing a two-color TSP method is also summarized.
Global versus Local -- Mach's Principle versus the Equivalence Principle
Singleton, Douglas
2016-01-01
The equivalence principle is the conceptual basis for general relativity. In contrast Mach's principle, although said to have been influential on Einstein in his formulation of general relativity, has not been shown to be central to the structure of general relativity. In this essay we suggest that the quantum effects of Hawking and Unruh radiation are a manifestation of a {\\it thermal} Mach's principle, where the local thermodynamic properties of the system are determined by the non-local structure of the quantum fields which determine the vacuum of a given spacetime. By comparing Hawking and Unruh temperatures for the same local acceleration we find a violation of the Einstein elevator version of the equivalence principle, which vanishes in the limit that the horizon is approached.
A Solar System Test of Mach's Principle with Gravimetric Data
Unzicker, A; Fabian, Karl; Unzicker, Alexander
2006-01-01
We present a new test for a possible Mach-Sciama dependence of the Gravitational constant G. According to Ernst Mach (1838-1916), the gravitational interaction depends on the distribution of masses in the universe. A corresponding hypothesis of Sciama (1953) on the gravitational constant, $c^2/G = \\sum m_i/r_i$, can be tested since the elliptic earth orbit should then cause minute annual variations in G. The test is performed by analyzing the gravity signals of a network of superconducting gravimeters (SG) which reach a precision of $10^{-10} m/s^2$. After reducing the signal by modelling tidal, meteorologic and geophysical effects, no significant evidence for the above dependence is found.
Quantum interference in an asymmetric Mach-Zehnder interferometer
Trenti, A.; Borghi, M.; Mancinelli, M.; Price, H. M.; Fontana, G.; Pavesi, L.
2016-08-01
A re-visitation of the well known free space Mach-Zehnder interferometer is reported here. The coexistence between one-photon and two-photons interference from collinear color entangled photon pairs is investigated. Thisarises from an arbitrarily small unbalance in the arm transmittance. The tuning of such asymmetry is reflected in dramatic changes in the coincidence detection, revealing beatings between one particle and two particle interference patterns. In particular, the role of the losses and of the intrinsic phase imperfectness of the lossy beamsplitter are explored in a single-port excited Mach-Zehnder interferometer. This configuration is especially useful for quantum optics on a chip, where the guiding geometry forces photons to travel in the same spatial mode.
Spatial heterodyne spectrometer based on the Mach-Zehnder interferometer
Cai, Qisheng; Xiangli, Bin; Du, Shusong
2015-11-01
Spatial heterodyne spectroscopy (SHS) is a new kind of Fourier-transform spectroscopic technique capable of very high spectral resolution. In this paper, a spatial heterodyne spectrometer based on the Mach-Zehnder interferometer (MZ-SHS) is proposed. It is modified by replacing one mirror in the Mach-Zehnder interferometer with a diffraction grating. This technique retains many of the advantages of traditional SHS. Moreover, the spatial frequency of the interferogram is strictly linear with wavenumber. We describe the concept of the new MZ-SHS and elaborate the exact expression of the interferogram. Also, a design example and two kinds of imitated interferograms are presented in this paper. One is simulated in MATLAB and the other is generated in ZEMAX using ray tracing method. The retrieved spectra from these two interferograms show a good agreement with the theoretical results.
Meyer, R. R., Jr.
1978-01-01
The static longitudinal and lateral directional characteristics of a 0.035 scale model of a first generation jet transport were obtained with and without upper winglets. The data were obtained for take off and landing configurations at a free stream Mach number of 0.30. The results generally indicated that upper winglets had favorable effects on the stability characteristics of the aircraft.
Emergent physics on Mach's principle and the rotating vacuum
Jannes, G
2015-01-01
Mach's principle applied to rotation can be correct if one takes into account the rotation of the quantum vacuum together with the Universe. Whether one can detect the rotation of the vacuum or not depends on its properties. If the vacuum is fully relativistic at all scales, Mach's principle should work and one cannot distinguish the rotation: in the rotating Universe+vacuum, the co-rotating bucket will have a flat surface (not concave). However, if there are "quantum gravity" effects which violate Lorentz invariance at high energy, then the rotation will become observable. This is demonstrated by analogy in condensed-matter systems, which consist of two subsystems: superfluid background (analog of vacuum) and "relativistic" excitations (analog of matter). For the low-energy (long-wavelength) observer the rotation of the vacuum is not observable. In the rotating frame, the "relativistic" quasiparticles feel the background as a Minkowski vacuum, i.e. they do not feel the rotation. Mach's idea of the relativity...
Three-dimensional flows in a transonic compressor rotor
Reid, Lonnie; Celestina, Mark L.; Dewitt, Kenneth; Keith, Theo
1991-01-01
This study involves an experimental and numerical investigation of the three-dimensional flows in a transonic compressor rotor. A variety of data which could be used, in a complementary fashion, to validate/calibrate the computational fluid dynamics turbomachinery code and improve understanding of the flow physics, were acquired. Detailed radial survey data which consisted of total pressure, total temperature, static pressure and flow angle were obtained at stations upstream and downstream of the rotor blade. Detailed velocity and turbulence profiles were obtained upstream of the rotor and used as the upstream boundary conditions for the numerical analysis. Calibrated flush-mounted hot film probes were used to measure wall shear stress on the hub and casing walls upstream of the rotor. The blade-to-blade shear-stress angle distributions were obtained at two axial locations on the rotor casing, using flush-mounted hot film probes. A numerical analysis conducted using a three-dimensional Navier-Stokes code was compared with the experimental results.
Spectral variability in transonic discs around black holes
Zampieri, L; Szuszkiewicz, E
2001-01-01
Transonic discs with accretion rates relevant to intrinsically bright Galactic X-ray sources ($L\\approx 10^{38}$-$10^{39} {\\rm erg s}^{-1}$) exhibit a time dependent cyclic behaviour due to the onset of a thermal instability driven by radiation pressure. In this paper we calculate radiation spectra emitted from thermally-unstable discs to provide detailed theoretical predictions for observationally relevant quantities. The emergent spectrum has been obtained by solving self-consistently the vertical structure and radiative transfer in the disc atmosphere. We focus on four particular stages of the disc evolution, the maximal evacuation stage and three intermediate stages during the replenishment phase. The disc is found to undergo rather dramatic spectral changes during the evolution, emitting mainly in the 1-10 keV band during outburst and in the 0.1-1 keV band off-outburst. Local spectra, although different in shape from a blackbody at the disc effective temperature, may be characterized in terms of a harden...
Preliminary measurements of aerodynamic damping of a transonic compressor rotor
Crawley, E. F.; Kerrebrock, J. L.; Dugundji, J.
1980-01-01
The aeroelastic behavior of a transonic compressor rotor operated in the MIT Blowdown Compressor Facility has been examined by means of piezoelectric motion sensors at the base of each of the 23 blades. Excitation has been observed due to rotating stall, due to an incipient flutter, and due to the facility startup transient. A method has been found for determining the aerodynamic damping force by modal analysis of the blade motion. Application of this technique to the example of excitation by rotating stall has led to the conclusions that the blade loading decreases in the stall cell, and that the damping force on the blades in the clean flow is in phase with blade velocity but opposite it in sign, leading to a logarithmic decrement of 0.2. This method of force derivation has quite general applicability as it requires only blade motion data such as are routinely acquired with strain gages. It is argued that models are needed for aerodynamic damping which focus on the effects of near neighbors of a given blade, since flutter often results in large response of isolated blades or small groups of blades.
Performance of turbulence models for transonic flows in a diffuser
Liu, Yangwei; Wu, Jianuo; Lu, Lipeng
2016-09-01
Eight turbulence models frequently used in aerodynamics have been employed in the detailed numerical investigations for transonic flows in the Sajben diffuser, to assess the predictive capabilities of the turbulence models for shock wave/turbulent boundary layer interactions (SWTBLI) in internal flows. The eight turbulence models include: the Spalart-Allmaras model, the standard k - 𝜀 model, the RNG k - 𝜀 model, the realizable k - 𝜀 model, the standard k - ω model, the SST k - ω model, the v2¯ - f model and the Reynolds stress model. The performance of the different turbulence models adopted has been systematically assessed by comparing the numerical results with the available experimental data. The comparisons show that the predictive performance becomes worse as the shock wave becomes stronger. The v2¯ - f model and the SST k - ω model perform much better than other models, and the SST k - ω model predicts a little better than the v2¯ - f model for pressure on walls and velocity profile, whereas the v2¯ - f model predicts a little better than the SST k - ω model for separation location, reattachment location and separation length for strong shock case.
A Detailed Investigation of Staged Normal Injection into a Mach 2 Flow
Eklund, Dean R.; Northam, G. Burton; Hartfield, Roy J., Jr.
1990-01-01
A study of the staged injection of two jets of air behind a rearward facing step into a Mach 2 flow was performed using the SPARK 3-D Navier-Stokes code. Calculated mole fraction distributions were compared with an extensive set of planar mole fraction measurements made with a laser induced iodine fluorescence technique. A statistical measure, the standard deviation, was used to help assess agreement between calculation and experiment. Overall, good agreement was found between calculated and measured values. Generally, agreement was better in the far field of the injectors. The effect of grid resolution was investigated by calculating solutions on grids of 60,000, 200,000, and 450,000 points. Differences in the solutions on the two finer grids were small. However, the mole fraction distributions were distinguishable. The effect of turbulence modeling was investigated by employing three different algebraic models for the jet turbulence: the Baldwin-Lomax model, the Prandtl mixing length model, and the Eggers mixing length model. Overall, the Eggers mixing length model was found to be superior for this case. Finally, the effect of the jet exit conditions was examined. A recently proposed Mach number distribution at the jet exit was found to slightly improve agreement between measurement and calculation.
Hypersonic characteristics of an advanced aerospace plane at Mach 20.3
Mccandless, R. S.
1985-01-01
Wind-tunnel studies have been performed in the Langley Hypersonic Helium Tunnel Facility to obtain static longitudinal and lateral-directional aerodynamic characteristics of an advanced aerospace plane concept. The nominal test conditions are a Mach number of 20.3 and a Reynolds number of 6.8 x 10 to the 6th power per foot at angles of attack from 0 to 25 deg and angles of sideslip of -3 and 0 deg. Stability and control characteristics are obtained for several deflections of the elevators, elevons, and rudder. In addition, a modified canopy is examined. The results indicate that this vehicle is longitudinally stable at angles of attack near the maximum lift-drag ratio. Also, the vehicle is shown to be directionally unstable with positive dihedral effect.
Yaros, S. F.; Carlson, J. R.; Chandrasekaran, B.
1986-01-01
An effort has been undertaken at the NASA Langley Research Center to assess the capabilities of available computational methods for use in propulsion integration design studies of transonic transport aircraft, particularly of pylon/nacelle combinations which exhibit essentially no interference drag. The three computer codes selected represent state-of-the-art computational methods for analyzing complex configurations at subsonic and transonic flight conditions. These are: EULER, a finitie volume solution of the Euler equation; VSAERO, a panel solution of the Laplace equation; and PPW, a finite difference solution of the small disturbance transonic equations. In general, all three codes have certain capabilities that allow them to be of some value in predicting the flows about transport configurations, but all have limitations. Until more accurate methods are available, careful application and interpretation of the results of these codes are needed.
Yaros, Steven F.; Carlson, John R.; Chandrasekaran, Balasubramanyan
1986-01-01
An effort has been undertaken at the NASA Langley Research Center to assess the capabilities of available computational methods for use in propulsion integration design studies of transonic transport aircraft, particularly of pylon/nacelle combinations which exhibit essentially no interference drag. The three computer codes selected represent state-of-the-art computational methods for analyzing complex configurations at subsonic and transonic flight conditions. These are: EULER, a finite volume solution of the Euler equation; VSAERO, a panel solution of the Laplace equation; and PPW, a finite difference solution of the small disturbance transonic equations. In general, all three codes have certain capabilities that allow them to be of some value in predicting the flows about transport configurations, but all have limitations. Until more accurate methods are available, careful application and interpretation of the results of these codes are needed.
Nonlinear dynamics approach of modeling the bifurcation for aircraft wing flutter in transonic speed
DEFF Research Database (Denmark)
Matsushita, Hiroshi; Miyata, T.; Christiansen, Lasse Engbo
2002-01-01
The procedure of obtaining the two-degrees-of-freedom, finite dimensional. nonlinear mathematical model. which models the nonlinear features of aircraft flutter in transonic speed is reported. The model enables to explain every feature of the transonic flutter data of the wind tunnel tests...... conducted at National Aerospace Laboratory in Japan for a high aspect ratio wing. It explains the nonlinear features of the transonic flutter such as the subcritical Hopf bifurcation of a limit cycle oscillation (LCO), a saddle-node bifurcation, and an unstable limit cycle as well as a normal (linear......) flutter condition with its linear pan. At a final procedure of improve a quantitative matching with the test data. the continuation method for analyzing the bifurcation is extensively used....
The use of hot-wire anemometry in transonic periodic flow
Bodapati, S.; Lee, C.-S.
1984-01-01
The unsteady wake profiles of an airfoil with an oscillating flap were measured in the NASA Ames 11 x 11-foot transonic wind tunnel. Laser Doppler Velocimetry (LDV) and holography techniques were used in limited region where optical accessability is available. X-hot-film wire was used to measure the wake profiles in the complete region to obtain magnitude and direction of the flow. A thorough calibration was carried out to determine the sensitivity coefficients of the hot-wire in three different tunnels at transonic speeds. A calculation procedure is established to resolve the hot-wire signals at transonic speeds and applied in the measurements of steady and periodic wake profiles. The effect of flow incidence on the hot-wire signals is evaluated and incorporated in the analyses. Typical hot-wire results are compared with the results of LDV, holography and pitot-static tube embedded with Kulite transducers.
Transonic solutions of isothermal galactic winds in a cold dark matter halo
Tsuchiya, Masami; Nitta, Shin-ya
2013-01-01
We study fundamental properties of steady, spherically symmetric, isothermal galactic outflow in appropriate gravitational potential models. We aim at constructing a universal scale free theory not only for galactic winds, but also for winds from clusters/groups of galaxies. In particular, we consider effects of mass-density distribution on the formation of transonic galactic outflows under several models of the density distribution profile predicted by cosmological simulations of structure formation based on the cold dark matter (CDM) scenario. In this study, we have clarified that there exists two types of transonic solutions: outflows from the central region and from distant region with a finite radius, depending upon the density distribution of the system. The system with sufficiently steep density gradient at the center is allowed to have the transonic outflows from the center. The resultant criterion intriguingly indicates that the density gradient at the center must be steeper than that of the predicti...
CFD Predictions for Transonic Performance of the ERA Hybrid Wing-Body Configuration
Deere, Karen A.; Luckring, James M.; McMillin, S. Naomi; Flamm, Jeffrey D.; Roman, Dino
2016-01-01
A computational study was performed for a Hybrid Wing Body configuration that was focused at transonic cruise performance conditions. In the absence of experimental data, two fully independent computational fluid dynamics analyses were conducted to add confidence to the estimated transonic performance predictions. The primary analysis was performed by Boeing with the structured overset-mesh code OVERFLOW. The secondary analysis was performed by NASA Langley Research Center with the unstructured-mesh code USM3D. Both analyses were performed at full-scale flight conditions and included three configurations customary to drag buildup and interference analysis: a powered complete configuration, the configuration with the nacelle/pylon removed, and the powered nacelle in isolation. The results in this paper are focused primarily on transonic performance up to cruise and through drag rise. Comparisons between the CFD results were very good despite some minor geometric differences in the two analyses.
Polytropic transonic galactic outflows in a dark matter halo with a central black hole
Igarashi, Asuka; Mori, Masao; Nitta, Shin-ya
2017-09-01
Polytropic transonic solutions of spherically symmetric and steady galactic winds in the gravitational potential of a dark matter halo (DMH) with a supermassive black hole (SMBH) are studied. The solutions are classified in terms of their topological features, and the gravitational potential of the SMBH adds a new branch to the transonic solutions generated by the gravity of the DMH. The topological types of the transonic solutions depend on the mass distribution, the amount of supplied energy, the polytropic index γ and the slope α of the DMH mass distribution. When α becomes larger than a critical value αc, the transonic solution types change dramatically. Further, our model predicts that it is possible for a slowly accelerating outflow to exist, even in quiescent galaxies with small γ. This slowly accelerating outflow differs from those considered in many of the previous studies focusing on supersonic outflows in active star-forming galaxies. In addition, our model indicates that outflows in active star-forming galaxies have only one transonic point in the inner region (∼0.01 kpc). The locus of this transonic point does not strongly depend on γ. We apply the polytropic model incorporating mass flux supplied by stellar components to the Sombrero galaxy, and conclude that it can reproduce the observed gas density and the temperature distribution well. This result differs significantly from the isothermal model, which requires an unrealistically large mass flux. Thus, we conclude that the polytropic model is more realistic than the isothermal model, and that the Sombrero galaxy can have a slowly accelerating outflow.
Lo, Kin Hing; Zare-Behtash, Hossein; Kontis, Konstantinos
2016-09-01
Flow separation control over a three-dimensional contour bump using jet in a Mach 1.9 supersonic free-stream has been experimentally investigated using a transonic/supersonic wind tunnel. Jet total pressure in the range of 0-4 bar was blowing at the valley of the contour bump. Schlieren photography, surface oil flow visualisation and particle image velocimetry measurements were employed for flow visualisation and diagnostics. Experimental results show that blowing jet at the valley of the contour bump can hinder the formation and distort the spanwise vortices. The blowing jet can also reduce the extent of flow separation appears downstream of the bump crest. It was observed that this approach of flow control is more effective when high jet total pressure is employed. It is believed that a pressure gradient is generated as a result of the interaction between the flow downstream of the bump crest and the jet induced shock leads to the downwards flow motion around the bump valley.
A least squares finite element scheme for transonic flow around harmonically oscillating airfoils
Cox, C. L.; Fix, G. J.; Gunzburger, M. D.
1983-01-01
The present investigation shows that a finite element scheme with a weighted least squares variational principle is applicable to the problem of transonic flow around a harmonically oscillating airfoil. For the flat plate case, numerical results compare favorably with the exact solution. The obtained numerical results for the transonic problem, for which an exact solution is not known, have the characteristics of known experimental results. It is demonstrated that the performance of the employed numerical method is independent of equation type (elliptic or hyperbolic) and frequency. The weighted least squares principle allows the appropriate modeling of singularities, which such a modeling of singularities is not possible with normal least squares.
Optimum design of camber line shape in a transonic axial fan cascade with splitter
Institute of Scientific and Technical Information of China (English)
SUN Xi-miao; WANG Song-tao; JIANG Bin; QIANG Xiao-qing; WANG Zhong-qi
2008-01-01
Genetic algorithms and artificial neural networks method were adopted to .optimize the camber line shape of splitter cascade in a transonic axial fan. For a certain type of transonic axial fan, the isentropic effi-ciency was improved by 0.6% with the modification of splitter cascade rotor. Numerical simulation studies show that the curvature of the optimum splitter cascade with larger suction side is increased, the static pressure gradi-ent is reduced, the process of flow expansion on suction surface is controlled and the separation near hub is e-liminated by changing the shock wave structure. The efficiency near hub is improved by 1.3%.
Institute of Scientific and Technical Information of China (English)
无
2000-01-01
Based on the standark κ-ε turbulence model,a new compressible κ-ε model considering the pressure expansion influence due to the compressibility of fluid is developed and aplied to the simulation of 3D transonic turbulent flows in a nozzle and a cascade.The Reynolds averaged N-S equations in generalized curvilinear coordinates are solved with implementation of the new model,the high resolution TVD scheme is used to discretize the convective terms.The numerical results show that the compressible κ-ε odel behaves well in the simulation of transonic internal turbulent flows.
CONDUCTION IN LOW MACH NUMBER FLOWS. I. LINEAR AND WEAKLY NONLINEAR REGIMES
Energy Technology Data Exchange (ETDEWEB)
Lecoanet, Daniel [Department of Astronomy and Theoretical Astrophysics Center, University of California, Berkeley, CA 94720 (United States); Brown, Benjamin P.; Zweibel, Ellen G.; Burns, Keaton J.; Oishi, Jeffrey S. [Kavli Institute for Theoretical Physics, University of California, Santa Barbara, CA 93106 (United States); Vasil, Geoffrey M., E-mail: dlecoanet@berkeley.edu [School of Mathematics and Statistics, University of Sydney, NSW 2006 (Australia)
2014-12-20
Thermal conduction is an important energy transfer and damping mechanism in astrophysical flows. Fourier's law, in which the heat flux is proportional to the negative temperature gradient, leading to temperature diffusion, is a well-known empirical model of thermal conduction. However, entropy diffusion has emerged as an alternative thermal conduction model, despite not ensuring the monotonicity of entropy. This paper investigates the differences between temperature and entropy diffusion for both linear internal gravity waves and weakly nonlinear convection. In addition to simulating the two thermal conduction models with the fully compressible Navier-Stokes equations, we also study their effects in the reduced ''soundproof'' anelastic and pseudoincompressible (PI) equations. We find that in the linear and weakly nonlinear regime, temperature and entropy diffusion give quantitatively similar results, although there are some larger errors in the PI equations with temperature diffusion due to inaccuracies in the equation of state. Extrapolating our weakly nonlinear results, we speculate that differences between temperature and entropy diffusion might become more important for strongly turbulent convection.
Calibration and Performance of the AEDC/VKF Tunnel C, Mach Number 4, Aerothermal Wind Tunnel
1982-06-01
Beattie - Bridgeman equation of state for air is shown in the figures included in this appendix. Real-Gas Enthalpy General Forms The following... Beattie - Bridgeman equation of state for air. 106 AEDC-TR-82-6 P(Rea1 Gas) R(P/PT)(P/PT)Idea1 8 PT1. 02 Sym 0 Computed Real-Gas Values Curve Fit TT, oR ~600...chamber properties; then the results were adjusted to include the real-gas effects. The real-gas properties are based on the Beattie - Bridgeman equation
Performance Limiting Flow Processes in High-State Loading High-Mach Number Compressors
2008-03-13
stage matching, and thus the performance of such machines. As such, the understanding, empiricism , and guidelines which apply well to machines of lower...discrete vortex with opposite circulation to the previous one is shed. A vortex street which is " locked " to the rotor passing is thus formed downstream of...255-6802 x231 (email: [)ouglas.Rabc,(wpatb.af.rnil) Dr. John Adamczyk, retired scientist from NASA GRC, has also contributed much to the research 18. 0
Numerical prediction of flow induced noise in free jets of high Mach numbers
Schönrock, Olaf
2009-01-01
A direct aeroacoustic simulation methodology is developed on the basis of the numerical schemes implemented in the commercial tool ANSYS CFX. The focus lies upon the efficient and direct numerical prediction of the flow-induced noise generated by natural gas and pneumatic applications. The respective compressed gas related components are characterized by tiny supersonic gas jets, strong noise emissions, poor accessibility by measurement techniques and excessive simulation costs in particular...
Numerical Simulations of Flow in a 3-D Supersonic Intake at High Mach Numbers
Directory of Open Access Journals (Sweden)
R. Sivakumar
2006-10-01
Full Text Available Numerical simulations of the compressible, 3-D non reacting flow in the engine inlet sectionof a concept hypersonic air-breathing vehicle are presented. These simulations have been carriedout using FLUENT. For all the results reported, the mesh has been refined to achieve areaaveragedwall y+ about 105. Mass flow rate through the intake and stagnation pressure recoveryare used to compare the performance at various angles of attack. The calculations are able topredict the mode of air-intake operation (critical and subcritical for different angles of attack.Flow distortion at the intake for various angles of attack is also calculated and discussed. Thenumerical results are validated by simulating the flow through a 2-D mixed compression hypersonicintake model and comparing with the experimental data.
High Mach-number collisionless shock driven by a laser with an external magnetic field
Directory of Open Access Journals (Sweden)
Morita T.
2013-11-01
Full Text Available Collisionless shocks are produced in counter-streaming plasmas with an external magnetic field. The shocks are generated due to an electrostatic field generated in counter-streaming laser-irradiated plasmas, as reported previously in a series of experiments without an external magnetic field [T. Morita et al., Phys. Plasmas, 17, 122702 (2010, Kuramitsu et al., Phys. Rev. Lett., 106, 175002 (2011] via laser-irradiation of a double-CH-foil target. A magnetic field is applied to the region between two foils by putting an electro-magnet (∼10 T perpendicular to the direction of plasma expansion. The generated shocks show different characteristics later in time (t > 20ns.
Boundary Layer Trip Performance Test on a 7-deg Cone Model at Mach Number 8
1983-10-01
b r a t i o n Re ldenha in Rotary Encoder ROD700 Resolut lo=:O, O000 des Dvora l l Accuracy: 0-001 des P a n a x e t r i c a MG-IOI Mois...LAYER STABILITY TEST PUN NUHuER 1028 PAGE 1 / - % DATE COMPUTED I1-0Cm~83 DATE R~CORDEO 2-~1 TI~E RECORDED 71~56~0 TIHK COMPUTED 09130 PROJECT... kiln NJOm;~Vle Jo~k LRETA 1.053E*03 1.821E+O$ 2.192E*03 2.553E*03 3.104E+03 4.54bE+03 6 .352E*03 80101E*03 10068E*04 1.250E*04 1.52~Et04
DEFF Research Database (Denmark)
Pradera-Mallabiabarrena, Ainara; Jacobsen, Finn; Svendsen, Christian
2013-01-01
-compact surfaces are involved. Here the generation of noise is dominated by the interaction of the flow with a surface whose maximum dimension is shorter than the wavelength of interest. The analysis is based on the surface-source term of the Ffowcs Williams-Hawkings equation. The acoustic source data of the flow...
Comparison of Experiment and Analysis for a High Primary Mach Number Ejector
1977-05-01
ure the secondary total pressure, also recorded on the HP plotter. A 30" (76.2cm) mercury manometer was used to measure directly the secondary total...supply pressure were readjusted to give the required total secondary pressure reading on the mercury manometer . Heat was added to keep the air streams at
Influence of Mach Number and Dynamic Pressure on Cavity Tones and Freedrop Trajectories
2014-03-27
1 0 ) ; 157 A. p12= p o l y f i t ( x ’ ,A. fcon , 1 2 ) ; 158 A. p14= p o l y f i t ( x ’ ,A. fcon , 1 4 ) ; 159 A. p16 = p o l y f i t ( x ’ ,A...x ) ; 164 A. f i t 1 2 = p o l y v a l (A. p12 , x ) ; 165 A. f i t 1 4 = p o l y v a l (A. p14 , x ) ; 166 A. f i t 1 6 = p o l y v a l (A. p16
Asymptotic Steady State Solution to a Bow Shock with an Infinite Mach Number
Yalinewich, Almog
2015-01-01
The problem of a cold gas flowing past a stationary object is considered. It is shown that at large distances from the obstacle the shock front forms a parabolic solid of revolution. The interior of the shock front is obtained by solution of the hydrodynamic equations in parabolic coordinates. The results are verified with a hydrodynamic simulation. The drag force and expected spectra are calculated for such shock, both in case of an optically thin and thick media. Finally, relations to astrophysical bow shocks and other analytic works on oblique shocks are discussed.
High-Speed Noninvasive Multi-Parameter Laser Diagnostics for High-Mach-Number Flows Project
National Aeronautics and Space Administration — Numerous ground test and wind tunnel facilities are used extensively to generate forces and moments as well as surface measurements of test articles required to...
Growth of a gas bubble in a supersaturated and slightly compressible liquid at low Mach number
Mohammadein, S. A.; Mohamed, K. G.
2011-12-01
In this paper, the growth of a gas bubble in a supersaturated and slightly compressible liquid is discussed. The mathematical model is solved analytically by using the modified Plesset and Zwick method. The growth process is affected by: sonic speed in the liquid, polytropic exponent, diffusion coefficient, initial concentration difference, surface tension, viscosity, adjustment factor and void fraction. The famous formula of Plesset and Zwick is produced as a special case of the result at some values of the adjustment factor. Moreover, the resultant formula is implemented to the case of the growth of underwater gas bubble.
Guilloud, G.; Schram, C.; Golliard, J.
2009-01-01
Despite the aeroacoustic expertise reached nowadays in air and ground transportation, energy sector or domestic appliances, reaching a decibel accuracy of an acoustic prediction for industrial cases is still challenging. Strong investments are made nowadays by oil and gas companies to determine and
Weatherill, W. H.; Sebastian, J. D.; Ehlers, F. E.
1977-01-01
A finite difference method for solving the unsteady flow about harmonically oscillating wings is investigated. The procedure is based on separating the velocity potential into steady and unsteady parts and linearing the resulting unsteady differential equation for small disturbances. Solutions are obtained using relaxation procedures. The means for improving the solution stability characteristics of the relaxation process are explored. A direct procedure is formulated which permits obtaining solutions for combinations of Mach number and reduced frequency for which the relaxation process has proved unstable. The pressure distribution for an aspect ratio 5 rectangular wing oscillating in pitch is presented.
Hot-wire calibration in subsonic/transonic flow regimes
Nagabushana, K. A.; Ash, Robert L.
1995-01-01
A different approach for calibrating hot-wires, which simplifies the calibration procedure and reduces the tunnel run-time by an order of magnitude was sought. In general, it is accepted that the directly measurable quantities in any flow are velocity, density, and total temperature. Very few facilities have the capability of varying the total temperature over an adequate range. However, if the overheat temperature parameter, a(sub w), is used to calibrate the hot-wire then the directly measurable quantity, voltage, will be a function of the flow variables and the overheat parameter i.e., E = f(u,p,a(sub w), T(sub w)) where a(sub w) will contain the needed total temperature information. In this report, various methods of evaluating sensitivities with different dependent and independent variables to calibrate a 3-Wire hot-wire probe using a constant temperature anemometer (CTA) in subsonic/transonic flow regimes is presented. The advantage of using a(sub w) as the independent variable instead of total temperature, t(sub o), or overheat temperature parameter, tau, is that while running a calibration test it is not necessary to know the recovery factor, the coefficients in a wire resistance to temperature relationship for a given probe. It was deduced that the method employing the relationship E = f (u,p,a(sub w)) should result in the most accurate calibration of hot wire probes. Any other method would require additional measurements. Also this method will allow calibration and determination of accurate temperature fluctuation information even in atmospheric wind tunnels where there is no ability to obtain any temperature sensitivity information at present. This technique greatly simplifies the calibration process for hot-wires, provides the required calibration information needed in obtaining temperature fluctuations, and reduces both the tunnel run-time and the test matrix required to calibrate hotwires. Some of the results using the above techniques are presented
Recent developments of axial flow compressors under transonic flow conditions
Srinivas, G.; Raghunandana, K.; Satish Shenoy, B.
2017-05-01
The objective of this paper is to give a holistic view of the most advanced technology and procedures that are practiced in the field of turbomachinery design. Compressor flow solver is the turbulence model used in the CFD to solve viscous problems. The popular techniques like Jameson’s rotated difference scheme was used to solve potential flow equation in transonic condition for two dimensional aero foils and later three dimensional wings. The gradient base method is also a popular method especially for compressor blade shape optimization. Various other types of optimization techniques available are Evolutionary algorithms (EAs) and Response surface methodology (RSM). It is observed that in order to improve compressor flow solver and to get agreeable results careful attention need to be paid towards viscous relations, grid resolution, turbulent modeling and artificial viscosity, in CFD. The advanced techniques like Jameson’s rotated difference had most substantial impact on wing design and aero foil. For compressor blade shape optimization, Evolutionary algorithm is quite simple than gradient based technique because it can solve the parameters simultaneously by searching from multiple points in the given design space. Response surface methodology (RSM) is a method basically used to design empirical models of the response that were observed and to study systematically the experimental data. This methodology analyses the correct relationship between expected responses (output) and design variables (input). RSM solves the function systematically in a series of mathematical and statistical processes. For turbomachinery blade optimization recently RSM has been implemented successfully. The well-designed high performance axial flow compressors finds its application in any air-breathing jet engines.
Impinging Jet Resonant Modes at Mach 1.5
Davis, Timothy
2013-01-01
High speed impinging jets have been the focus of several studies owing to their practical application and resonance dominated flow-field. The current study focuses on the identification and visualization of the resonant modes at certain critical impingement heights for a Mach 1.5 normally impinging jet. These modes are associated with high amplitude, discrete peaks in the power spectra and can be identified as having either axisymmetric or azimuthal modes. Their visualization is accomplished through phase-locked Schlieren imaging and fast-response pressure sensitive paint (PC-PSP) applied to the ground plane.
Temperature sensitivity of waveguide Mach-Zehnder interferometer
Sokolov, Viktor
2013-01-01
This thesis is part of a project that aims to develop a sensor for the detection of methane in the air and in water based on a waveguide Mach-Zehnder interferometer. The main application of this sensor is monitoring the environment and the ability to detect a leakage of methane. The development of a sensor includes analysis of operational conditions. In this project one of the greatest concerns is temperature. The temperature difference can reach several tens of degrees in the air, and severa...
Quantum logic processor: Implementation with electronic Mach-Zehnder interferometer
Sarkar, Angik; Bhattacharyya, T. K.; Patwardhan, Ajay
2006-05-01
An approach for implementation of quantum logic in electronic Mach-Zehnder interferometer (MZI) has been described in this letter. All single qubit gates can be achieved by electron spin manipulation using Rashba spin-orbit coupling. Double qubit gates can also be implemented using the orbital degree of freedom of the electron. The MZI can be realized with intertwined ballistic nanowires. Spin injection and detection in the system can be done by a mesoscopic Stern-Gerlach apparatus. The system can be coupled in an array to form the quantum logic processor.
On Mach's Principle and the "Special" Theory of Relativity
Ashura, Uzumaki
2016-01-01
First, we present a history of the school of thought that the Cosmic Microwave Background Radiation acts as an ether in language familiar to high school students in English-speaking countries. Then we illustrate the properties of this ether and of a hypothetical "test mass" using a brand new thought experiment. Finally, we recount some post-Einstein efforts at a mathematical formulation of Mach's principle and raise some questions about what implications it has for the locality of rotation and for quantum gravity. This paper does not prove Einstein wrong.
Directory of Open Access Journals (Sweden)
Cui Michael M.
2005-01-01
Full Text Available To reduce vibration and noise level, the impeller and diffuser blade numbers inside an industrial compressor are typically chosen without common divisors. The shapes of volutes or collectors in these compressors are also not axis-symmetric. When impeller blades pass these asymmetric structures, the flow field in the compressor is time-dependent and three-dimensional. To obtain a fundamental physical understanding of these three-dimensional unsteady flow fields and assess their impact on the compressor performance, the flow field inside the compressors needs to be studied as a whole to include asymmetric and unsteady interaction between the compressor components. In the current study, a unified three-dimensional numerical model was built for a transonic centrifugal compressor including impeller, diffusers, and volute. HFC 134a was used as the working fluid. The thermodynamic and transport properties of the refrigerant gas were modeled by the Martin-Hou equation of state and power laws, respectively. The three-dimensional unsteady flow field was simulated with a Navier-Stokes solver using the k−ϵ turbulent model. The overall performance parameters are obtained by integrating the field quantities. Both the unsteady flow field and the overall performance are analyzed comparatively for each component. The compressor was tested in a water chiller system instrumented to obtain both the overall performance data and local flow-field quantities. The experimental and numerical results agree well. The correlation between the overall compressor performance and local flow-field quantities is defined. The methodology developed and data obtained in these studies can be applied to the centrifugal compressor design and optimization.
Calibration of the 7—Equation Transition Model for High Reynolds Flows at Low Mach
Colonia, S.; Leble, V.; Steijl, R.; Barakos, G.
2016-09-01
The numerical simulation of flows over large-scale wind turbine blades without considering the transition from laminar to fully turbulent flow may result in incorrect estimates of the blade loads and performance. Thanks to its relative simplicity and promising results, the Local-Correlation based Transition Modelling concept represents a valid way to include transitional effects into practical CFD simulations. However, the model involves coefficients that need tuning. In this paper, the γ—equation transition model is assessed and calibrated, for a wide range of Reynolds numbers at low Mach, as needed for wind turbine applications. An aerofoil is used to evaluate the original model and calibrate it; while a large scale wind turbine blade is employed to show that the calibrated model can lead to reliable solutions for complex three-dimensional flows. The calibrated model shows promising results for both two-dimensional and three-dimensional flows, even if cross-flow instabilities are neglected.
Numerical computation of transonic flows by finite-element and finite-difference methods
Hafez, M. M.; Wellford, L. C.; Merkle, C. L.; Murman, E. M.
1978-01-01
Studies on applications of the finite element approach to transonic flow calculations are reported. Different discretization techniques of the differential equations and boundary conditions are compared. Finite element analogs of Murman's mixed type finite difference operators for small disturbance formulations were constructed and the time dependent approach (using finite differences in time and finite elements in space) was examined.
Wai, J. C.; Sun, C. C.; Yoshihara, H.
1982-01-01
A method using a transonic small disturbance code with successive line over-relaxation is described for treating wing/fuselage configurations with a nacelle/pylon/powered jet. Examples illustrating its use for the NASA transport research model are given. Reasonable test/theory comparisons were obtained.
Transonic solutions of a wing/pylon/finned store using hybrid domain decomposition
Newman, James C., III; Baysal, Oktay
1992-01-01
Transonic Euler calculations about a complex multicomponent configuration are presented. The 3D Euler equations are solved utilizing an upwind-biased, alternating direction implicit, approximately factored, multigrid algorithm. Computational results are compared to experimental data of the finned store in a carriage position.
Modelling flow phenomena in time dependent store release from transonic aircraft
CSIR Research Space (South Africa)
MacLucas, David A
2014-07-01
Full Text Available Computational Fluid Dynamics is routinely used in clearance of stores for carriage and release from aircraft in the transonic range of flight. A well-known validation case is modelled in this study, for which aerodynamic loads have been compared...
The Application of the Probabilistic Collocation Method to a Transonic Axial Flow Compressor
Loeven, G.J.A.; Bijl, H.
2010-01-01
In this paper the Probabilistic Collocation method is used for uncertainty quantification of operational uncertainties in a transonic axial flow compressor (i.e. NASA Rotor 37). Compressor rotors are components of a gas turbine that are highly sensitive to operational and geometrical uncertainties.
Metrology Measurements of the DSTO Transonic Wind Tunnel Store Support Arm
2013-12-01
UNCLASSIFIED UNCLASSIFIED Metrology Measurements of the DSTO Transonic Wind Tunnel Store Support Arm Adam Blandford, John Clayton...provide quality assurance for test clients. This document details metrology measurements that were conducted during February and March 2013 on the store...Australia 2013 AR-015-818 December 2013 APPROVED FOR PUBLIC RELEASE UNCLASSIFIED UNCLASSIFIED Metrology Measurements of the DSTO
Wing-Alone Aerodynamic Characteristics to High Angles of Attack at Subsonic and Transonic Speeds.
1982-11-01
indicators of symmetry since the wings were unbanked within the limits of tolerances and flow angularity. Longitudinal, spanwise, and vertical... unbanked wings at subsonic and transonic speeds from low to high angles of attack. The wing planforms varied in aspect ratio and taper ratio with
Symmetries and Group-Invariant Solutions for Transonic Pressure-Gradient Equations
Institute of Scientific and Technical Information of China (English)
王丽真; 黄晴
2011-01-01
Lie symmetry group method is applied to study the transonic pressure-gradient equations in two-dimensional space. Its symmetry groups and corresponding optimal systems are determined, and several classes of irrotational groupinvariant solutions associated to the symmetries are obtained and special case of one-dimensional rarefaction wave is found.
Cosmological constant implementing Mach principle in general relativity
Namavarian, Nadereh; Farhoudi, Mehrdad
2016-10-01
We consider the fact that noticing on the operational meaning of the physical concepts played an impetus role in the appearance of general relativity (GR). Thus, we have paid more attention to the operational definition of the gravitational coupling constant in this theory as a dimensional constant which is gained through an experiment. However, as all available experiments just provide the value of this constant locally, this coupling constant can operationally be meaningful only in a local area. Regarding this point, to obtain an extension of GR for the large scale, we replace it by a conformal invariant model and then, reduce this model to a theory for the cosmological scale via breaking down the conformal symmetry through singling out a specific conformal frame which is characterized by the large scale characteristics of the universe. Finally, we come to the same field equations that historically were proposed by Einstein for the cosmological scale (GR plus the cosmological constant) as the result of his endeavor for making GR consistent with the Mach principle. However, we declare that the obtained field equations in this alternative approach do not carry the problem of the field equations proposed by Einstein for being consistent with Mach's principle (i.e., the existence of de Sitter solution), and can also be considered compatible with this principle in the Sciama view.